NACA - Industry Conference on Personal Aircraft Research

• / _TACA - L-_nDU_RY COI__.-_.EITCE A Conpi2abien of the by NACA Staff ._:_',pe._.'s !h,eccntcd l:e_f_erg Langley Memorial Aerr,nauti_'al Laborntory Lsn_e_ Field, _fa. Septer_er 20, 1946 // TABLE OF CONTENTS Pa_e INTRODUCTION LIST TECHNICAL FLY_NG QUALIT LES OF CONFEREES PAPERS . ................ .............. PRESENTED ...... •" • ' i 2 6 6 ....................... History _z_i Significance of Measured Flying Qualities . . by Melvin N. Cough ......... Flying Qu_lltlee Requirements by William for H. Personal Phillips ....... Airplanes STABILITY AND . . . C01_i_ROL 9 14 . ................... Effects cf Individual Stability Factors on Flying Characteristics of Airplanes . . . by Marion O. McKinr_y, Jr .............. Proportioning by Charles Design the Airplane for Lateral J. Donlan ................ St_rfaces Stability . . • 21 of Control . . • by Thomas A. Toll . . . 27 Prediction of Control-Surface Hlng_-Momeut Characteristics . • . by Charles W. Frick, SAFETY ANDSPINNING ..................... Jr ..... 33 38 A Flight Investigation to Increase Light Airplane . . . by Paul A. the Safety of a Hunc._r ....... . . . 39 Factors Affecting Spinning of Ligh_ Airplanes by Anshal I. Neihouse ................ 42 Development of Airfoils and Hi@h-Lift Devices by Laurence K. Loftin, Jr ................ Maximum Lift and Stalling . o . by Harold H. . . . 47 o . 51 _// Swoberg 77 81 the Structure and Statistics of 82 Atmospheric Gusts AS_rplane .. . Parkinson POWER PT_d_S Design .. Rubert . Wilson... . Aerodynamics by K_rm_dy 103 Icing of Engine Induction by Willson H... by Arthur ...PROPE_S .. .... of Power Plant Ins_llatlons F.. .... by Jo?m C... Systems ..... ..... Va....... ...ey Fiold. ... o .... ... ... ...... .._tions Aircraft D_si_ .... Office. by AIRC"RAFT Some LOADS Remarks .............. Crigler Propeller Noise ..... 122 The authors are mombers of the staffs of the NACA and Moff_t Luboratories Field. .. SEAPLANES • _ .mal by HarrAson C.... ...... ... Ohio. 72 76 Performance Gains D_sign .. 91 lO'_ . Maneuvering Loads on Vertical and Tail Surfac_s .. by in Relation to the Pers_>nal Charles C.... • • . on by Attention to Detail Herbert A.. ... .. ........... Propeller _]oloction an_ Effloienc_ by John L......... Hunter Prospects Porsonal _EW R_EARCH Cur!_ent an_ Proposod _CA of Inte!.. A.. ... ...... Noise .... by Henry A..... for ....... ...... . .............. Vogeley DRAG CLEAN-I_P . . Horizontal Pearson .... Reg!er 59 65 Problems in the Design of Propellers for Reduction . _ ........ _]anders ll7 121 Research Invustig. 86 90 Seaplane and Amphibian by Jo_m B..... ! ........ Chandler_ llO for Now Types of Proptulsive Sy:_t_ms Aircraf_ .... Jr........... Shufflebar_r . and the NACA Washington CJove!and. ......... at Lang_....s_ in Pers.... .. .. .. ... Jr ....... by Arthur W........e.......... Calif. !_RODUCTION This conference was organize_ by the NACA t) ac._uaint ai.rp.l.._ne designers with th,_ r_.;s_Itc ,.)f _artLme research considered t:) be. (:f _.nterest _n the design of personal aircrsft a.nd.to discus_ cvu'rent and propose& NAC_ research prcJects _pp]_icsbie t_:, light airp!anc..s. The technical papers pres,JJ_.tcd at the co.a.. ference were nece_Js_irily brief _mct ,._'c rcl.)r+duc,_d herein in the s_¢,c form _s they _¢oro 6;].vc,n in the _nterest of prompt distribu_ion. Thc_,_ ,_g',1-,_l rresentat__on ;rod this compilati,_u arc cons".,icred as complementary to, rather th._.,._ s_bstib_.it_s for, the CommitteeVs complete and formal reports. A number of reperts usod as references heroin, which were originally issued in classified form; are no_" declassified a_id are being printed and ,_istr_but,_d as NACA Wartime Reports. A list of the conferees is included, f. NATIONAL ADVIZ [)P,Y COMMIT_2 FOR #_'.}_OI,IATFflCS 2 LIST 0F _7_v_ The followin_ _mllere!Ei_'_tcredat tDe i.[ACA-_Tndustry Conference on Personal Aircr._ft Research, l,anglt;i _&.-m_ria]. Aeronautical Labcr.'_tcry, Len(Cl_y Field, V_,, oepte_,e. '_ li ": Abbott, Ira _. .fd_!er, L.P. _crustein_ Dr. Karl Baird, Robert Breder, Sam C. _ro_, Dayton T. B'_rn_ E .R. Burrows Ch_zndler, H.C.: Jr. Child. H. Lloyd Collins, Leighton Cooke, Richard Corddry, Charles Crigler; Jo}n_ L. Cr _wl._.y, J.W. Davis, D.J. Dearborn, C.H. &e _'lorez_ R.._dm, Della.yon: Hugh Diehl. Don!an, Dralcy, Dyccr_ Capt. W.S. Char_es J. E.C. C.F. Donal_l R. Ranger A!rcz'oft Enzinos Goodyear Ai_-craft Corp. Wa;o Aircraft C,.mpany £:2_n Aeronautical Ccrp. Or ;_'m A:i.r_=raft Coo#year Aircoolod I[ACA Ck, i Air F_cts MaCaz:'n,_ _la!]. Street Jou_.um! United Press NACA NACA Ly_cming NACA U.S .N. Kut.ional Goodyear BuAer NACA NAC'. CA _, A.M.C. Division, Aviation Corp. Aircraft Motors, Corp. Inc. r,uis P.esearch ._drcraft Cot_cll I/" Eastm:m_. MaJ. Eaton, Frsd Eiserer, !.on l_v_:_s George _ Flovor_ Frick_ Don C.W., Liaison 0ff!co I_ epubli c Aviation Am,orl can l'.vi,;_ ti on Ccm_ont_ea! th Aircraft Cessna NACA Aircr_ft Jr. C&r Cos , Don C_rteis, Jerry GCuting, Joseph Oi!ruth_ ]_.R. T. Fairchild Engine _md Airpl&ne Cessna. Aircraft Aircraft _udustrioc Assoc. NACA Corp. List of Co._er_es (cont '&.) A.M.C. NACA Liaison Offico Goddel!, _hj. H.C. C-ouch, Melvin N. C riswo!d: R.W.? I!. Gue!icD, I{.V. Hacks bhai. Paul l,udiuf:_en.-Griswold, inc. C, & A. Aircrofb, _c. Aerc_qtic A_-rcraft Propal3er_ De:_&ix k,riab_ on Corp. l,t].j._n_ _a _].ne _'D_, _.,_:_curAim-craft i]nr:in,_s Eepubl-c Aviation Cert. Ai_.-craft O_,_ers & _2i:]-o_:,s _saociation Ey'_n Aero_mut_ I_ACA Path Finder sal C.:rp. Ha&ley_ Josso Z sr.11injFred !:a:,u_n_n, T., Jr. Ha:rd_n, Oeorcs Hartrunft, J.B., Hayer, R il_h Hei__s _ C .H. Kess, !_arl Koekstz'a, H.D. Ho!co_o ilo_._ard: Hulbez't H_ter: Hunter, ; it. Je_,n Ross _ F,.A. Psul A. W.H. Jr. C c_dyear Lifo,raft Aircraft Industr_os Assoc. Cen_Inenta]. Notors Corp. NACA NACA kc lom_,:ll A!,'c_'_ft Co"!:. G. & ,_.._.ir'r::'f't,, Aircraz't ]mc. Ccrp. 0ffJ co "_'_ _ sbra_,mt _ E.U. Jr_mouneat b Walter A. _"iper Kermey,, _pt. Raym)na _ilpatrich, R.A. Kinkead 0 Robert K_usckj F.J. Kurt, FranP_in T. _est._z b Loening, Loftin_ E.H. Grover Laurence A.:.[.C. Liaisen i _" • _ _, "_,_'l _ &_rpo, " _, H_mpton _s.:Lianca Aircraft Or_:_x'_n Aircraft Rzn6;or &ircrsfb !_g. Engines Corp. 4"* K., Jr. Ma!oy, R.B. May, George McClax.rou_ R. McKi_u_,ey, M.O._ Jr. MoLto, Josoph _4cS_ro!y_ A!exandor Melquist, Victor Mockler, Don _. Mosebach. M.A. CAA Consolid:_tod-V'altoe N_CA NACA Air_raft Corp. Kaicor i:!c<ri_wings_ !_c. A T".at_on !7 _.:u ,'_ _ _S,.,OC Ai_'cr..,,_Industries _ A:h'cruft 3L_&ustrios Assoc. Oo,:_dyear Aircrafb • . Tb:_'.heed EACA I. A..J. Consoli .-_ . John C. H. A.H.I.u?..!.tcot_.nn._ otl_r u._storn N_wspap_r Uni._'-' c. _.C.Jin3a_i_: M. E. S!mc. _'.n IL"J".. l_e!-ry. Aircraft Taylor En.. Piper_ T. John Sweborc.... I. Sho:. Jo_m H.W.-_ ]grother s N{'_CA N\CA Aircraft Xnd'astric.i_.I_'_0 Cc. __": <"" Los!:!. ShPrp!ess. D... Sgi__uer: I:. T. Shsrp.d. C.':\CA I.-oarfircraft Aircraft C_mcrs . I<. Po_rsoT!.e!d.eCorp. Shuff!oborger. S..s • _. Aircraft _.C.F.Lir __ a_ _ Co.ingz. Corp..A i._r. Sanders.M._ on. Piper: _. U:.iACA l_.W..'.I. & A. St. So]iambs ck: O.G. _ '_" _"" C-. Aircraft [_chvoizo'r Aircraft I.. _. Inc. . R elrA Z I.". IiACA _[&'30 .-):'thr.: A=scc.L. NACA liper iJpor Aircraft Aircraft Corp.:<7-o_ c_'_:_tAviai-i':':_ COla._ . _Yei_ e_nyer_ Harry Eeihou_!e.. W. Times <3 /" So'l!e_ E:_rtley A.uLncer!n7 C_.rscn: E. I_ .ACA Luscomb_ N_n: York E. _a!ph .)l.P._.% Pilots Assoc.. _erk_nccm_ P<me.:io _. Sm/.'a or it Sh.C.ACA Alrp]-._ACA EACA Frood_man Aircraft _]ns. C.o C. Tale. D._.E. Robi... o!!dated-Vultec. E.__. _. Aircraft ..>rta!. S:mucnl cA B_'_.E... Dr. L!ai3(. s r_ "I] To%.T.vics Assoc. Harvey .F. JoD_u B.n¢iu_b. I:ACA I. _O- E.orl.A._atod-Vul.ircruft r'.CA IU-CA Loc!_. Inc . <iii _etor_on_ !var i_hiili_s. Ai_'_r_ft.K.hode.t_c. Corp.. Joseph A. _ e__. ... F. . _ a.V.th. Dr.Li_t of n_-_-_=n_ (cgnt '_ Scn_cni (._ Tu%ti('. R.n Off'i co Eomick_ Y:ubort.I'-CA C onr.%CA A .__ _r F. _ NACA I_. . 1.i_a'tinCom?an_ . LycJmins Division. Fred E.) Aer::m2.ctory_ John F. La_ence Wilson. Weick. Vo61e!ey. VJ. T. Wie2msm. McDonnell Airclaft Corp. Lee v_'_ go!m. Wri_t. C. A'Jiation Corp. NACA C_t Jr._esearch R Corp.e G!e ra'l L.P. Herbert A. _i!li_n_:s. Arthur W.. Wi!lem D.C£ Encineering _.t of Co_eerees Harold (oont'd.tic Aircraft Propellers Ulrich_ Van Atta. Int::rnatJ cnal News :_ervz .5 Li2.H. 6 FLYING QUALIT.TES . ludin_ five of th9 priv:%te-o_.re set of roqt_r<._t the bogim_ing of the war had _:ssemble& complete quantitatl.llities in terms of qtlantJtles that b.ss:_ry h_strumentation an(] the accumulation of specific d_:f._.airplane t_pe.tr_nent'_tion was apparent as more specific meas'_'ements of _..nel tests.. Th_-_re have been mm_y c_ntributors n.f ai_pl-_nes tested was incre?ose& to 44 at the Langley Lu.ndto 16 at the new Amcc Laboratory.co of flying om_d maneuvering an airplane.been entered into _:ntl. all NACA f'_c..t!c _uuntitat:_ve measurement of flying characteristics has._CE 0F _SURED FLYl_. N'._. the work of the Azray and Navy testing centers. interest in this directionwas _'Iso ovid. ._nts a_d res-_tln_ ai_lane mot..ions on existing end new military _rcraft._ntrol movem.ncrcasei need for rec<. of exlctinz des!zns and future developments.anes. The A_y and. ana .. reoe.ntinu_d to apply its faci!Itle.-.o:_ in[. in_.. pre&icted from win&-tum. and looked to the I<ACA t. inforn_. _:_iy i_ 1941: for the NACA to prei_.e war effort.c<. _% pilots alik_._stion acct_ntO.nd much logiest]_ deTelopmgnt.s possib!_. Gough The flying qualities of an airplane may be defined as the stability -and control characteristics that have _.. From the f'_nad oi' _nf¢_z.r <_.usiastica/ly by producers. M_m%facturcrs of milit_ory desiTns were f_. The endeaver to aevelop a systemt_.cte_ it wc. The J._tinu. All have approci_-_ted the necessity to evaluate the hitherto ur_atisfactory relative opinions of pilots regarding their f!i_t cxperiences. . By the end of the war the total nt_b_.ated in these t(. The hACA c<.i!it_s wore directed to the i_me@i_tt._.need by the (_ugge_/qcim Safe Airplane Competix. In the past. their i. engineers: o_erators._'eive airp].ons were required..ion.mem" cr flight. °.r(].. _qqcn this cour_try entered the war. th<.TG %UALITIES By Melvin N.tmiliarized with the info_unation then available..-._tlom on t_.m import:-_:_tbearing on the safet_ of flight and on the pilot's impressions of the an.monts a for seti_f_ctoryflying qu'. tho CA. Pilots themselves desire a more factual metho_ of expl'ession" and en_.urch programs of th_ NAC&.vy revised the general ?_ACk Fly_ng Qualities Specifications t.._medlahe specific requirements.v._ its investing'at.neers require quantitative informat"on for knowlcd_%.% _tfcty prozram.boz'_tor_ "_.d been measure& in fllg_ht and cotuld be calculated by cng_inecr_ and could ev_h be.HISTORY A_VD SIGNIFfC_._r_(!. s to the development cf the ncc_!_. : R_e. other epeskors will aspects of the existing discuss s(_e (_f the knowledge of flying REFERENCE 1. which_ when met. Even in the absence of quantitative measurements._-.e to be considered as a very minimt:m. and expan&_d as necessar. It and many other rep(_rts of s_ec_ific measu_'emonts m_.u better uudorst_diI_ of why an ai_Tlane h_.. The _\CA requirements_ as they now exist. leased for general us_._ systematic approach understandable to both pilot and engineer has been doveloood. NACA ACE. the ro]atlvo opinions valus. give reasonable ass1_-u_co that the airplane will be safe to fly and desirable from the pilot _s standpoint.¢hich must meet unusually high standa-. 8 This vast experience. R. It is of interest to note that not only does ther_ now exist . it will be desirable to obtain characteristics superior to those specified. which may be flo_ by unskilled pilots and _. April Sat] sfactory 194!. F!yinc. _ay.les :_s it do_s and what is be?. but a new tezmlino]og_. for GilrutLj R.ieve& desirable. is now available to all._ :_ :t_ . of pilcts familiar with the requirements are of greater Throughout the specific engineering qualities._ in the light of new &evelop'menbs and are now in the fo_m of _l I_ACA report (reference 1). The general NflCA specifications were contim_lly modified. . For a light airplane._s._._s cf flight safety. an& _. Q _.de on airplanes are to be re.@ .uirements of Air_:]an. . _.FLYIT$G ITTESREQ__f_ QU.-. Phillips . with the same degree of accuracy that it may be measured in flig/qt or in the _ind tur_nel. First take-off. both from flig$_t and wind-tunnel tests.-acteristics._pl% see references 3 and '.une. it is now possible to predict from the alrpl:me':.) The requirements for the longitudinal trimming d(_vice are on the list. a great deal of data have been acctunule.wL.' dimensions the t?ower-off static longitudinal stability of an airplane. for s_ti.ri. 1^rhile previously there was considerable speculation as to the flying characteristics _!es_rod in _ a_rp]. nco 2. In addition. are listed the requirements for the elevator control in Tests of ntnnerous wind-tunnel models have be_on made with ability of a gro_. These requirements were set up as a result of er_erience gained in testing all t. including five light airplanes.gfactory flying qualities of airpl_ules.ted on the effects of power on stability.&Y. e by the designer of personaS. The control forces re0uired to Cly st_adily at v_ious speeds may al_o be predicted-as a r_s'_lt of extensive NACA research on control-surface hinge mou:cnts _u_d aero<yn_nic b. A report is available which stm_n_%rizes the next . T'_ble I presents a list of the factors considered in the requirements for longitudinal stability and control characteristics.:_d _n referc. The requirements may be listed under the general headings of longitudinal stability and control ch_"acteristlcs. PLA_ES A great deal of r:_search work _as been conducted du_'ing the war yea-._Isnces. I _m_ild like to point out how NACA research has contributed to the km. l_teral stability and control cha.nts for flight._dcoof some of these topics._ONAL H. _nd stalling ch_Lracteristics. it is now possible to specify in quantitative tel_ns the minimum requirements that an airplsno must meet in order to be cons_de:-_d satisfactory from the pilot's standpoint. Flight-test data are _u_'_._pl_es elevator control in steady theft has bccn made of the (r_ference I). By William FO_.<!D. The requirements may therefore be used _th confidenc.d hence the basic elevator control characteristics in steady flight. As a rusult of an an_lysls flig_ht tests of n_. N_xt are the rc:.._rpes of airplanes. (For ex.md-board in place in the tunnel to determine the airplanes of various designs to meet this requircmc_nt.zaerous ai_..'s to determine requirements.qulrome. airplanes to smrive at a design with satisfactory stability and control characteristics. It. _e light airplane manufacturer should be encouraged to note that in the ca_3e of several fighter-type airplanes.lO re.i_[ted to definite quantitative values in the requirements. The characteristleo of the uncontrolled longitudinal motion are considered next. The requirements for aileron control characteristics are expressed quantitatively in te_ns of the minimumv_lue of the helix angle generated by the wing tip in a roll.e n_y fail to lJ_et the require_. Quantitative limits for the control motlons and force gradients necessary on airi_lanes of real4 v _ifferent types have been determined from flight measurements.ucted on this subject in the past therefore has little bearing on the subject of fly. Table I also showsthe requirements for lateral stability and control. 'l_neproblem of reducing the tr._nts. whore the problems of roclucin@ the control force changesdue to flaps and power are muchmore difficult than for a sI. or phu@_old oscillation of an airplane have been fouz_dto be rel_tlvely uuimportant. The elevator control characteristics in landing are specified in the final longitudinal requirement. This method has been ¢levelopodfrc m an analysis of the results of flight tests of n_. This r_quirement . _.vul_s of manywind-t_. A new airplane may be designed to incorporate these characteristics by the methods described previously. however._iningthe pilot's ooinion of an airplane.-_e-_-ous airplanes. This report (reference _) will allow senaccurate estimation of the tab sizes to meet this requirement° The elevator cont_'ol characteristics in accelerated flight have been fotuld to be very important in deter_._m changes in the case of a li_t airpl_ne should therefore be fairly _asy..or elevat_r eCfectiveness. Extensive wind-tum_el data are also available on this subject. Tr_m changes due to power and flaps are 1. Sometimesthe short-period-oscillation characteristics of an airpla_._ng qualities. may be of interest to _ntion in passing that the characteristics of the long-period._aller airplane.u_el ani flight tests on the effectiveness of trimming tabu. as well as theoretical analyses to showhowto avoid the undesirable characteristics (reference 7). designs with very small trim changes have been developed as a result of wind-tu_el and fli_-_ht tests. The elevator an_le required to land is of ini_ortanc_ because this is frequently t'_e most critical requirement f. Recent reports are available which describe flight experiences with uni_atisfactory characteristics (reference 6). A report is available which presents a methodfor calculating the elevator _ngle required to land (reference 8)._ extensive theoretical work cor_. Ordinarily changing . This requlrament has been fo_d to be a critical one from the st_qdpoint of directional stability._doqu:%tu directional cross-wind stability. The force characteristics and the pitching moment due to . A largo amou_u_ of direction_ul stability has n_ver failed to be beneficial to the h_uudling ch_.-.r ar_ ty.rious flight conditions. The _%_dder must per.. or_Ir_rily imogen as dlhodr::. doslgn does not greatly reduce the adverse yawin_ moments at hitch lift coofficients._m _%ny ftulcti_. providing satisfacto_'y control in _id_slips._t_ stability for flight with asymetric power._r. ado with alrpl:m.¸ The yaw due to ailerons should not exceed a certain m_ximum walue.ll was set uo before the war. as will be discussed in a l_%t_r p_per._o. The meet the trimming .tlons It.d by m_ny factors ether than the design of the _n_dd. %_e yawing moment due to sideslip or directional stability must b_ sufficient to meet certain requirements for providing satisfactory sidcsllp characterlstics_ limiting the sideslip in rolls due to aileron yawing moments.d next in the list of requi_'ements. 'rod providing adeq_.n_ besides sLmply providing directional trim in the vo. . Soy _ral flighb investig. _llo_ osti_ation of th_ t_%il size required to provi_c ._.cteristics of an airplane._s of various pl_ forms and different amounts of geomet_'ic dihedral have shown the allowable limits of dihedral effect for satisfactory flying qualities The • rudder control characteristics are consid.rudder and aileron trimming devices m_y be-de. A summary report on NACA lateral-control resgarch has been prepared (reference 9). and numerous subsequent tests have confirmed its validity. It has been shown by flight tests that the rudder _ff_:ctiredness required for take-off and la_ndi_g is detozmin. A great deal of rese:_rch has been done to dotermlne the _dder co_i_r. Th_so functions include offsetting the advo_-°s_ ail_[_on yawin Z moments._3ignod requiraments in the same man_er as the longitudinal devices. This report contains sufficient data to allow th_ d_sign of satisfactory ailerons f_._.he ailoroT.l off oct._e of airplane.:'.r_._r. providing satis__acto_y control dur_i_ag take-off and lar_d_Ing.o_t du:_ to asymotric power on _ multiengino ai_'pl_o. and offsetting th_ y:_twing mo_.ttions requirod to satisfy these r_quirements and in later lect_'es the design criterion for spin recovery will be discussed in mor_ detail. to The next r_quiremonts specify the limits of rolling zlomenL due to Sideslip. providing adequate spin recovery. Design dat_ ar_ availabl_ t.. e comparedto thu control friction so that the stick will have _( dci'inite c. 'lqqo_itch[nz momentdue to sldoslip should be small _.qv. The.7 with thu assur_mcc that its flying quaii tlcs will besat is factor 2 ._5 to a typical ll).trch ar. Th_ trim ch:_Z_L_ i0 due to changing flap or po_-ercondition should be v_ry _:_.z_ com_Lrcd to +fno friction t. Thu stall should be preceded by adcqu.importeaut.ed flight _hould be between 7 _.rkcd ]nc_. The st. there shoulO b_ __o rollir_ instabi]_. . The rcsulto of flight tests have been used to establish these roqulrements _mda grct_t deal of theoretical work performed by the NACAis available to enable the designer to predict these characterlstlc.12 sideslip.desi'p characteristics are shownin figtu'c !.:sc. [rcmonts for light a!rpl_e ut the NAC5 (reference In conclusion _t may be poi_uted out info_uuation has been accumul'vted both on s_:'_bisfP~ctor2" flying ouP_litic:_ and on th_ moans !'cr dcsl--gn_.. be superior to the minim_mrequirem'. The final roQu:irement for le. The characteristics of a personal airpl:_e should.cteristics measuredin steady sid_s!._n. ql_nijtles._rc very .ips. rud'l{r control should be m_fficlentl_ pow.'._ovide helix angle of the wing tip of greater ths_nC. _cn the _._n wJtn t.:..:d. Time does not permit a detailed discussion of all th_ requlrem_nts._peods.md should b_ sufficiertly lar_. Typical flight measurementsof s.. forr_ Lhe subject of the n_xt rogulroz_nt. A_-"_ rcsult_ _t is bo]._ flyin::._iieron control effectiveness s.-ign an airp]c_n6 of _h_ pcr_:'on::._me is sb_ll.ng_n airnlano to incorporate thes._odlflcution.l-_'_irp:'-<.houid be ._in_ in the fo_m of buff_tin_ of the c. Someof the cheracterlstics that are believed very desirable in a light airplane as a result of N_CA r.o p=.ty in any flight cond.o _ive the control stick a def'inito centering teni.07 r_dicn.raL.'..[on.:'. in rearward motion in the stick emd pu: i force .irpl-'..irnl.hc _tick held full back.noend by a m:.'. The force variation _. beth of which are char_. but _ufficiontlj lar.icved posoible to deP.t_ral stability and control is concei'nodwith ghu uncontro31_d lateral and directional motion. that a large amo_mb o_' the r_.nts for _atisf:_ctory flying qualities.ilezon yaw.::ncy.:_ as follows: the st:_tic longitudinal stability in elk flight conditions should be large_ so that t2e pilot will be warned of the approach to the stall by the rearwJ::_rd position of the cti._nd po_is per g. This charactur'stic was obfx%_n_d by rclaLiv_ ij minor .'_._ (ref<.rith accsl_ration in acceh. . The stalling charactcristlcs of a Torson_l-typ< a.::sthe stall is approsched.tc wuz. cw._n_fflcien_ [.ii._neral. q'n_ d_r_ctional stabi!it_" should be large. The _.i t.-:rp3':._+.ck-for(:e gradients in stralghb flight should b_ stable _. in-g_.renco lO). a TS_e _-_iloronforces _hou!d be Right.s this factor may l_ad tc Inadw.rtont stalling.isfactery s._-.:_rful to overcom_ a.:ntcrlng t._ndcncj.nt catcccr.ck at low fli<:ht . _. and Ste_fie!¢. and Murrej. ". Thomas A. Harry. May 19 _. tlie E].m L-<_._.. ) 1O. of St.lO_ _" Phillips: i. P._illiam H : A Flight _ "] Longitudinal Oscillations of _n Airplcne uith Short Peidod Free Elevator. lOb9. March 194_. $. Maurice D. R. : An:ilysis of of Tabs on C _ntro!-Surface 1946. Hunter. NACA AILR. Goranson.aue'J as Determined in Flig/st. 1944. J.Ovcrhcmg "__id F_i. _': Thomas 0. k_01.e. 1_i_te.%tudinol Stabil_ty o[" A . : Effect of Po'. . pl. NACA k_ig._nai_: & T}. 2ichard I. on the Effects NACA TN No. ._t a Light Airpls_no._mic C-. System. NACA ACE No. : Wind-Tlmnel Characteristics of Airplane 1943. R. Stewart Available Data Hinge Moments. N. N_utral Po!nts 2._ A.. Of M. Harry A. Cranda!l.kD. Toll): (Prospective _.e_-at.T 3 >_IDS._ IIvegt.an_s with F_:ee Ccn.. Data on the Aez'o<_vn.%CA AB_ No. !944. Greenberg..) I].I!H01.ntrol Surf acorn. : _mm/ysis c. Ptu-ser.n: DeflectS.:)r L4114. : A Fli_. ITACA CB No. A. Seers.Ri%Tlb_/_Ir_ES _ _n_te.. NACA ACR No.. m_<! Ste_mfie]d.trol_ Including EfCe_t of Fricti'-n in Control. 191_4.on Required 9 / A Method for Pre:licbind to Land. G_eenberg. Increase the Safety of . N:_:('!A _-14_ No. NACA C1_ No. Febi<. Z. 4.ieoret_cal InvestiGation of Len. Pa_l E. and Toll. 19!C.13.L_m-_aryof Research Department Lateral-Control (Ccnpiled Research.._er on the Stick--Fixed Neutr21 Feints of Several Single-Engine Monop]. D_ota on Control Surfaces B_l_ces. _'_d Vensel.f Av-_ilable Having Plain .ick-F_xed L%C01.l. Inv:_sticat_on (P_'ospoct_ve to _T{. Ear-v. l.9_'4. Leonard: A Theoretical Investigation of the Lateral Oscillations ef _n Air_l'ane wfth Free Rudder with Special Reference to the Effect of Frlct_on. Mau-rice D. : Estimation of Airplanes.. ]. Rudder-control F.OUTLINE OF FLYING QUALITIES REQUIREMENTS I. Uncontrolled lateral CONNITTEE FOR AERONAUTICS iii. Requirements control in landing Stability and Control for Lateral A. Elevator B. Cross-wlnd characteristics due to sideslip force characteristics due to sideslip and directional NATIONAL motion ADVISOI_Y <) H. Stalling Characteristics . Yawing moment G. Uncontrolled longitudinal F. Requirements A. Limits of trim change due to power and flaps G. Elevator Ii. Rudder and aileron trimming devices D.0q O_ °r-4 _z Table I . Limits of rolling moment due to sideslip E. Yaw due to ailerons C. Aileron-control characteristics B. Elevator for Longitudinal J Stability and Control: control control in takeoff in steady flight <if> C. Longitudinal D. Elevator trimming device in accelerated motion flight control E. Pitching moment I. 0 PUSH LEFT I0030 LEFT \ _ELEVATOR \.- Steady sideslip characteristics.20UP RIGHT IOCONTROL POSITIONS AND BANK 0 ANGLE. I0DOWN LEFT ELEVATOR ....-/" TO_TA_L ___ _ _--_/ .(/'I_ _. LB. DEG.. DEG. NATIONAL COMMITTEE ADVISORY FOR AERONAUTICS .L-ANGLE "_/" OF BANK N "_ 20-1 IO0PULL RIGHT CONTROL FORGES.uooER 1 20 30 RIGHT Figure I. \ t ! I 20 I0 0 SIDESLIP ANGLE. l I0 t-. 14 STABILITY AND CONTROL I . ters are _:ccoml_lished by vatting the center-of-gravity pos. the lateral d f]_yJng quali+.llt_e.i_s w_ ll bc second.lan.ng-w. For the sake.__r. ON CHARACTERISTICS By Marion O. He.o d_+e_no the effects of these two factors on the lon_f tudinal flying qualities (reference I) are presented in f_ure I.il_n_Lh._t'.on.q qualities requirements may be divided into three _reuLcs . poor. fair._.s c_ncern_d x.-_. those concerned with the control cha. vity location ThQ_ center-of-g_vity location is presented margin which Ss the distance _n chords that is ahead of the neutral point or aerod_namic a_lane.locity factor I_Cm/O \ - chang<-: of ._.nter-of-gzB. The NACA has conducted some research during the _ar to determine.there it has been found convenient an_ cconomlcal to _ary the basic stability factors inderendentl_y of each oth.15 E3q_E. The demp_ng in pitch is presented _erivat_w_ coefficient Cn_ with which _s the rate of dangerous) arc shown as and damplnLg -In _.unncl +.:?in_ in Ditch or horizonta].itch.s treated first am.. The present papur _lll discuss the: effects of the individual stability factoz's on flying qua.-fli_[. In will be tl'eated the conventional manner th_ longitudinal flyJnr= nual[ti_.. horizontal or ver+_c__l tail size or +all length.ilot's opinions of the longitudinal st.re th. in te_ms of th_ st_._ith the stability of the airp. The _ro_or design of the controls and considerations of stalling and splr_:In_ will be discussed in some of the following papers._act_ristlcs of the airplane.r to bo s_fe and easy to fly.. Such variations of the individual stabil_. Phillips presented a d-[scussion of the flying qualities that an airplane should have in ord_:._. of convenience these flyin._'dral ancle._t +. and functions of the ce. Much of this work has been done _n the free-flight tunnel .those concerned with the stability of the airplane.t._. C For many 2e8rs the longitudinal stability has bo_:n cons_.lc the c_nter of gLcsv'tv center of the cometicto in tc_m_ of the stability '_ qc_ e pitch]. the effects of many of the basic stability fsct_rs on flying qualities. McKinney._CTSOF FLYING INDIVIDUAL STABILITY OF FACTORS AIP_ _IANES Jr. l The results of an investiqation in the fr.n'ed to be primarily dependent upon the center-of-cravtty location and the da_.'_ad_ne_s (as _ndicated by the: qualitative ratings: good. _a'l slz_ and '. and the dih. The previous paper by Mr..tTr paz_me.d. and those conc_o_mr:d with stalling and spinning._. o -*_ pilot thought. The effective dihedral is expressed in terms of the stability factor CZB \_j angle _hlch of i_ the rate and of the change range of rolllng-momcnt of values effectlvo of C_. of the wing ar_a located one-half of the wing span aft of the center of grav_t.e two parameters (dihed_ml effect.y. is presented dihedral.lity factors on the lateral flyin_ qualities.red angles ( with s_deslip.yand effcetlve directional stability is expresced in terms of the f.itudinal steadiness will be 6. _-.ocd..mter-of-g_av_ty posit_on. approximately dihedral It is apparent that the best lateral flight behavior dihed_. It is ajparent that the pilot of the model coDsidered tha+ the long_._ _:m]_.:. reference 2.ral an.!)Cnlwhioh CnG k _ ] with angle ch_rt is th_ rate. pendent upon th._:lc and vertical-tail in figure 2.ght behavior as functions of the did'cottons.onal of change of stability the derivative yawing-moment of Cn_ coeffi¢[ent covered and.G cocffic_.r_..y) .._alltics because the:_.-w_ng taille_s ty!2e to one bering a horizon+al tall 2J_ percent of the "_in7 ar_a 2 chords a_ _ of th_ center _'_ of gravi+y.cti_.08 cho_.tud_. Similar investigations have been made to dete_"m!ne <._l and hig_h directional It _as most _" . repre s _. The a_ea on the _ateral flying q_m]ii.n .._ c. The results of an investigation in the freeflight tunnel to determine the effects of diheS._ _ * vertical tails from approximately 0 to 30 percent. ies are presented This chart was prepared from data presented in The _ilot's opinion of the lateral fli.nal steadiness is virtually independent of the damg_ng in pitch and is almost entirely de.n. chart presented a pict_re had the effective believed _ that this of th(._ns]. that. The cPart indicates that if the cen_er of &_ravi-ly Ks more than 0.st_b:ilit.h_ effects of various basic lateral s+ab_. roblem p _n obtaining good lateral flyinq q. on the c_hart represent from -lO ° to 20 o.1 stabil:'_t. This conclusion has been supported by full-scale flight t_.sts.nt cov::.16 The range of rakes of Cmq covered on this chart represent a .'d _head of the aerod_u_antic center of' the airplane that the lon. and d_rc. % range of ai_lane configurations from the st_a__. +he model when it had a s_dl positive stability. the range of values on the of s_dezl_p. nondimens_. :n .ntwith the full-scale flight t.b_ _-ho NACA _n a light sirplane sever'_O_ years a t_o (reference 7).\ are the primary factors affecting lateral stability.litlcs investigations of quite a few full-scale J_.n_.rb.002 on directional which current stability corresponding to a value t:._.ral the flj_ng charactoristlcs w. With 9 ° diho_._. With the smallest dihedral angle the airpl.t'. As the dihedral t_._mpin_ _n' y_.m<_ters concerned with the lateral area and d. capubl_ ()f flying itself to the greatest possible extent.'-_.i/ Figure 2 has been correlated wi!_h th. however_ to haw spiral stability in the high-speed and cruising conditions and still have _ood lateral flig)_t beh_t_ior -_f the airplane.ocations off that ai.zht-h:tnd branch of the good flight behavior boundary of figu.c}vu't _n the as the dihedral :_nglo was varied from 0 e to o/> to 6" . qualities requirements d. mass distribution.. it will In order have to good laterc._ qualitD_'[_. 3 to 6.tllthan represents a som:_wht_.'._r. Tt is possible.'_tisfactory size vertic_. effects ef incro'_. That is.._rplan.o not specify that an airplane should be spira_.l tall is to be used _n a.1 owner alrplane so that it might be completely stable .r rei_on oC thSs cha. From these considerations it may be concluded that if the minimum ._nd 9 ° .OOl) of' Cno snd of 0._me was i'o_d to fly the best.ly stable. \.._results of f]yin_ que. u_d the aerodyn_ic p.._ = \ -O.u__nd hL]S boon found to be in good a_rcem_[.so and still have good lateral flight bc. it is realized. The: f..yin_. however.zht behav.tl stability might be a desirable feattu'e for a persem..t!argor is con('ra?. The spiral stability problem of light ai_q01anes was therefcre snalyzed. _ chart (fig. t:tre spe.%rid. the flying ch. that spir_.irplane should have about 5 ° effective dihedral (C7.ur s.osts. were so designed that it would fall near the ri._Jrplane.:s.. l_--V_t_igations were made. The resttlts of these investig_ations are prc_ented in referenc_.w of . In brief_ these investigations effect of reasonable changes boundaries of figure 2..?.s_ln6 he t dihedral illustrate angle we may refer to some flig_ht tests _d_ L.1 the f]._bo!s _n th.mdaries presented in figure 2.re quite .lY used li_oht airplan.. to determine how variatiol_s of certain other lateral stability parameters would affect the f!i. The factors inw_sti_4_bed were wing loading._r_ctorfst_cs became worse._gle was increased. It was found that it w_s not possible to have spiral stability over the er._'p_". 2) represent the l.havior. showed in these that there w'_s virtual ]y no factors on the fli_ht behavior <) Although the flying-."e '.:.lor bo_.d ran. if the effective d_hedral and directional st:tbil:!ty of an airplane would place it in the good flight behavLio.. The center of gravity must be far enou_ ahead of the aero_.iti_c._ prc_perly proportioned to obtain good l_tera! flying qualities. tail It may therefore be concluded _-ot_l_d have the better response smoother thro1_h a series that the s. SYMBOLS C_ rate of change of pitching-me.._ moment (suc_ as might be oroduced by a rolling ghost or by an aileron giving a pure rolling moment) are shown as function of time for an airp1_ie _¢ith two vertical Cn6 tails .ve satis_'actory flying qua]. that an'airplane have cert._ w_locities following an abr_pt rollin. 1 second and holds that rolling velocity fairly stea_ily. thus caucing the and s_me _irplane the effects of ch_ng.. Cn_ with = 0. .. It s_ms_ at present.a small The tail.. never reached a steady rolling state. hewever.d llq_ _ qualities. of rolling There ar_ other f:_ctors which affect the f. and the dihedral angle and _ertical tail ar_o_ L_ust b.bout the magnitude as these caused by to be very dlffic1_It to fly._ve not been considered in the present paper. _" such as the contz'ol and stalling characteristics which h.lz'pl_e with the larger to its controls and .18 poor.ler tail.. It is necem_ary. d p_ and dt _ _ were / greater for the airplane with the s_!ler tail._c center . its = O. The initial adverse yawing was greater for the airpla_e with the smaller tail: and the angular accelerations (_s _ndicated by the slopes of these curves. airplane maximum rolling velocity in about._ rolling aileron deflection which caused rolling moments _._ou!d ride gusts.yn_. primarily because of the adverse yawing duo t. The airplsa_e with the mm:O. merit cooffic.tion of three factors. 00168.f :.ir. that the problem of obtaining these basic stability char.tile vertical _ tail area figure 3 was prepared._n_.:in basic stability ch_r_:_ct_ristics in order to h_.00042: the larger and tail a larger reaches tail.nt with pitchir_ veloci _y factor . irpl_ne to obt_In _Tc:od longitudinal flying qualltios. however.0f the e. In o_ler to illustrate the ailerons._cterist ics c_u be reduced to a considers. The calm_lated roiling and 2awin.qc .. . feet b S M N L q V / _ring span.. yawing time.:le of I _cn_ C r_te of ch_go of rolling-moment coefficient with angle of sideslip ( _Cz _ -g2 Cm pitching-moment coefficient (. pitching feet foot-pounds respect to stability _._t.IIL . with foot-pom]ds pounds second raglans radia_ns per oct secc_n_. moment t 6 M L degr(_os pitching rcllin_%-moment with respect to stabil. feet per P r roll_in._t-p(_mds moment.s of sideslip. fo.F___ _ \qcS/ Cn __winz-mement coefficient \_2_ Cl roliin_-moment coefficient (_L).:_ { . yawir_g moment roll ing moment._x:s_ foot-?ounds . an_le velocity.s. .coeff-lcient with an_.i9 Cn_ rate of change of ya_. CL C lift :ne_ coefficient-\ aerodj_amlc feet square q_ )_-chord.rlng-m_ment .v . wi_'. second per square foot pressed'e.g _rea.x:_. second._ velocity. d_amic ail_pee&. e]. Jchn P. end Gough. Charles L . 1946.!0_.1 Ste. . }[ACA AT_ T{o. as Doterr_ined by Tests of % Mcd. ]91#6. : The Effect of Lc.bi!it[: __d Con+re _' • .. Sou]e_ Kartle_ A.1 in tLe Ereo-Fli_._'_ Dihed-.2O EE.nd Contrc.FERE2_CFm i.ral Stability ".t Turmel. 5. Weic__<. !{ACA ARR No._e Cha_-.n end Rot_-tionc._ _ k_._m.er_ with Diffcren¢ A_.5L05... 1094.i. Mc!vln 17. Jr.o_." F£fcct of WJ___ Loed. 4<)4•.ilmtlon of the 4 ] Effects of Dihe&ral_ Vert_. C'-_um_Dbe!l.ec%eric_tios of en ... l Dei_ing in Pitch on the Lonzitu#dnal Stability Characterlst-cs of an Airpl_ne as ]3etertuincd by Tests of a Model in the NACA Free-Flight TtuLnel. Dl_l. {F25_ 1943. am Stability an& Control Characteristics._':.:e_ H_. Charles L. L4FOe.t Wide Aicrons and Various Spoi]. ]..stigation of the L_te_al Control Chare_cto_Ist.focrt M. 1. ping .y T._. I'_ACA Rep • No._e& by Te-_ts of s. 4.. J_h. John P an@ Seacord. 3H3 I.: i. D_'ake.s_ Distl-lbution on the Leter. 1.teral Area o_ the _ • '" Stability and Control Chal-acter'_tlcs of &n Airplane us Dcte_i_....]._n Yaw on Late_ . in the Langley Freo-F!_ht Turn_el. i. 1944.-'_ 6. in the l'_reo-F!igkt Ttn:nel.ca_. _ The Effects of St':tic on_foe"_ _ John P o_d Pa1_isen.9. ct_n..:.: invc.. Jr.l_ r_: m • '1 Are_. }_ACA A[!?_No. NACA AP_ No. 2.._ W.. I_ACA TN No. I_ .: 194'_._al. 1943._... ic_ of [_. " A 1.:ctu_tsf < _..brg_..Mo4_].ne_ _%nd Altitude on La_.934 " . McKinnev_ Marion 0. and Lift Coefficient on Late-_-a! St_bi!ity az_d Control Characteristics. Jr • The Effect of _4. Fro_ E.-_.. : ]_. NACA TN N. _ubort M. _atur.. an& Seater6. 3.xoerlmental Determination of the Effc_cts of . Csm7 :be!l.'l" _ht 7._ Stability an& Rotter.: Experimenbal Detezzu. l Ch_ractcristlcs of on Airpi_no a_ "i'_bel_l -_Od by Tests of a Mode.. .- bCm bC L .004" .- Effect of vertical-tail area and dihedral behavior.STATIC (b MARGIN.oo% NATIONAL 1 ADVISORY .04L.-4 o . DIRECTIONAL STABILITY 1 in pitch F-22 FLIGHT TEST . \ \\ \\\\ \\\\ \ \ \ \ \\\ \\\\\\\ \\\\\ POOR \ \ \\ \\\ \\\ \\\k \\ \\\\\\\\\\\\\ DANGEROUS I I I I 0 4 8 12 PI TCH.C1B .oo2 0 I I I I I -.12- GOOD .Cm q 16 DAMPING-INFigure I.002 DIHEDRAL._.004 FiTure 2.08- \\\\ \ \ \\\\\\\\\\\\ \ \ \\\\\\\\\ \\\ FAIR .- Effect of center-of-gravity location and dampin_ of lon_itudinal steadiness. .16- .00"2 0 EFFECTIVE .f . on later:llflight . .3 .2 ANGULAR VELOCITY YAWl N G _ 0 I I I I I I I .00042 VELOCITY ANGULAR .I Cn 8 =0.C) 01#---0. .00200 C/. 6 7 NATIONAL ADVISORY COMMITTEE FOI_AERONAUTICS to a rolling Figure 3.I I I I I I I I On 8 = 0 . SEC. 00168 0 I 2 :3 4 5 TI ME.2I• 0 -.- Effect of directional stability on response moment. ve of rolling-moment coefficient against sideslip. however._hininghi_)1 values of Cn6 and low values of C_ have been amply iemonstrated.18n STABIT_ITY In proportioning the aiz_plane for lateral stability.PLANEFOR LAT_2AL By Charles J. We have found that side force _ove]. defined as the slope of the cu_. Methods for doing this are dlscusm_i extensively in papers by Pass (reference 3) _nd Murray (reference h) .) The side force developed by the vertic_. (See reference _-. and tLe effective dihed_:al factor C_.ly from design charts such as those presented in s. The contribution of the vertical tail to directional stability is proportional to the slope of the tail-normal-force curve. the propeller side force presents no sei'ious desi_ probl_m i!_a_nuch as the variation of the propelle_" side force _ith an_%e of sideslip can be predicted very _eliably and rapi(:'. In any event. (See reference lo) The unstable yawing moment of the fuselage also can be estimated fairly reliably either from experimental results on similar shaped bodies or by theories such as Mttlthopp:s. (2) the basic fuselage yawing moment NF. C(_nsoquontly._tio of Cno to CZ8 :_nd to determine how both of these parnme+. The purpose of this brief paper is to summarize the design factors that influence these parameters and to indicate how reliably they can be estlmated _ The basic forces influencing the directional stability of an airplane are indicated in figure 1. that the horizontal tall and fuse!e4e act as en endplate _ich may cause the effective aspect ratio of the vertical tail to differ from its geometric w_lue by 50 percent or more.rs and their ratio are apt to change for the various flight conditions encountered. however..oped by the propeller becomes important for !arze propell(_rs located at some distsnce from the center of gravity. it is rather important to estln_te ho_ large ths endplate effect m_J be. Don. The desi_ problem is to provide the optimum r_. which._ltail and its variation with angle of sideslip however cannot te estimated as reliably or as easily. and (3) the _Ide force on the vertical tall LT.paper by Ribner. __q_e advantages of _=c. is primarily a function Gf the aspect ratio of the tail. It must be remembered. The forces and moments contributing to the yawing moment about the center of gravity are: (1) the propeller side force N. in turn. the t_ characteristics of most imports_rice_re tile &iroctional stability factor Cna defined as the sSope of the curve of yawSn_-moment coefficient against sideslip.9ROP01_TI01TL_G TH!] AIi2. The t_pics-! effects produced by a dorsal fin are illustrated in figure _. These results _e_'e t.ting the directional stability. _e have f_equently foun_ _t necessary to increase the effectiveness of the vertical tall at large _ngles of sideslip. In general.nce5 _nd are for e_n aspectratio 6 wing with no gocmetric dihedral.?J_enf_°'cm refer(_.anesis. The relative wing position is represented along the horizont_l axis._ill notice that whereas the vertical tail _Ithout the dorsal fin becomes ineffective at om angle of sideslip of about 19°.s of estimating this effect together with the complications Intro<uced by flaps and po-_er are discussed extensively in several of the papers now available to you. The magnitude of the interf_rence effect is illustrated in figure 3. One device co_uonly empioyod for this purpose is the dorsal fin. The asymmetry of the curves is caused by the propeller slipstream and is typical of the effects introduced by power. You .p_reater dihedral effect than lo_-wi_Z a'_raugements. For high-po_ered airpls_cs. We have found that with low-wing arrangements that the _ihedral effect is reduced when power is applied. The si_e_ash effect is important and should be t_ken into acco_u_t in estim_. Metho_. We have fotm_l that the _i__ew_sh is generally favorable for lo_:-wing airpla_nes and adverse f_r high-wing airpL_-nes. The dihedral effect exhibited by a_l-p]. the ve_'tical tall in combinat_. In a_ditlon to supplying adequate stability at small angles of sldeslip._l tail i_ slso i_tfluenced by the side_ash associated with the flo_ about the fuselage-wing combination. of co_rse. the power effect on dihedral can be extremely pronounced vhen flaps are immersed v . but the effect _epends ta a consilerable extent "bn the amount of wing irmmersed in the slipstream. closely associated _Ith _ng position. high-_rlng arrangements exhibit _. Any dihedral effect introduced by setting the ring at a given geometric dihedral angle can be added to the ba_ arrangement. The fuselage cross-sectional shape appears to be of minor importance. but fairly reliable estimates c_n be made of the effect._o. The ordinate on the left is the effective f" -> dihedral parameter CI_ and the equivalent geometric dihedral angle is given on the ri_t. In this figure the yawing-moment coefficient Cn is plotted as a funCtion of the sideslip angle _ for an ail_lane vlth a vertical tail used in conjunction _Ith a dorsal fin. It _ill be noted that the high-wing position is equivalent to about 9o of effective dihedral and the low-wing position about .on with the dorsal fin retains _ts effectiveness out to 30° _thout _ny detrimental effects at small angles of sideslip.22 The effectiveness of the vertic. pounds i. on vertical force_ fig. N coefficient coefficient with propeller side force.down condition.!d_slip angle pressur3.nt of rolling-mcment coeff.ip s] ope of curve sidesl_p yawing-moment rolling-_oment aga__nst " Cn C7. foot-pot_'Ids rol]. potu_. SYMBO_T q Cn_ Ct[ slope of curve of ya_-ir_-mom_.ng area. airplanes. po_mds feet squsro feet potu_ds per degrees degrees square foot w_. b S q w_ng span. flap .23 _n the effect slipstream and.) Y T total th_Ist (S_e _._cient coefficient against side_]._ d_hedra] in the full-powcr. dynamic s. For low-powered the effect is probably _m. angle. foot-po_mds pounds coefficient with respect L to stab:il_ty X axis. yuwing moment r_!_sDect to stability Z axis.ing-moment foot-potmds basic side fuselage force sld_ yawing moment. of ba_k.rln_ lift.ds: also. many sere!co types e}_hibit negati_ _. . tail.import_t. NACA _ No. Herbert S. NACA Rep. 705.!..: Wind_. NACA A_l_ No. H. No. Pass. 775. 1946. NACA _i$[ No. Earry E. Ribner. m_d Vertical Fins. @md a Circular Fuselage. : _. ]0 _: 3. : _Totes cn the Propellel_ and _. 1940. and W_llac_.llpst_eam lu Relation to Stability.2... 2 Multhopp. _ Elliptical 012_. 5 Kouse. L4-_-!2a. / .Tu. H.__th Available Theory. R_Lfus O.:ind-T_mnel Investigation of Effect of Interference on Lateral-Stability Characteristic_ of Four NACA o _ Win_s. 4. 1941. R. : An_lysis of Wind-Tunuel Data on Directional Stability an([Control.19 4'_. _ti:Ltr R. : 19_. Murray. Ae__odynsmics of the Vuselege. 1050. NACA TN No..nne! Investi_atlon of End-Plate Effects of Horizontal Tails on a Vertic_l Tall Compared _. _.04-..- Effect of dorsal fin on directional stability...02 o \ l I I I I I _ I -04• / I 40 :30 20 !0 0 /8 -10-20-50-40 NATIONAL ADVISORY CONNITTE[ FOR AERONAUTICS Fibre 2../_i:_ -._.- Basic forces on an airplane in a sideslip.02-.r-'d © Y T WIND D I. 011 DORSAL FI N. .% fWITHOUT DORSAL FIN DORSAL FIN . WITH ._.._.wSIN Figure I. ADVISORY FOR AERONAUTICS ...Effect POSITION on dihedral NATIONAL COMMITTEE of wing position effect..ll... WING FiTure ( i 3.---4 o O [] ROUND FUSELAGE FUSELAGE EFFECTIVE DIHEDRAL -+5 ° ELLIPTICAL .0 ! LOW MID . H IGH . d bo rup_.> for changes in spued binations of the from the trim cond. Figur._s of bhesc t_. the hinge-moment characteristics of a given control s_face may be specified by two paran_ters . va!u_s of _Ch/_k_ -_ro given on the ordinate scale a_. Th._foru._d balances.ments has b_en at least as importa:_t as the problem of controLsurface effectJ.ng thes_ _ars we have been printa_:il:/ interested in the development of control surfac_s capable of C _ pr. Control c::aract_risti_=s for m_u%y flight con.nd for s_mfaces with welldesig_. SLni_ACES Toll From the year 194l to the present tim_. The greater part of recent r_search effort in control-surface desiC:_ therefore has been directed toward the general problem of hinge mgm.nt par_amotor._Ch/6m.. Values of the hinge-moment coefficient ar_ plotted against control-surfac_ deflection for various fixed angles of at_ck._iv_!u. w]. Figure certain be o_:. For most plain (or unbalanced) surfaces e. which is th_ variation of hing_o-moment coeff_cienc with ang_!_ (:f attack and :_Ch/OS.ued by the NACA labor_.venes._namic balance. A large part _o_.-i th_ of control following d{scussion is concerned with the application of results of recent research on hinge mon_nts to the design surfaces for light aircraft.%i_iens ma3 matically in b_. tl_er'._ch relate_ to contl_ol-surface design has been perfor_.ors. the problem of control-surface hinge m:. In this f!_.'itl: deflection.'._3._nts _u_.vo this boundary wc_Ald give a _tabl_ variation of _t_. a considerable a_uount ._/ru.ol flight.-._ r_sea. consequently.e<l.._ hlnge-mom_-. methods of ob taining ae_"_._o p-_ramet. Dur'l.i$ion hin(_e-mo_cnt p_._s of _Ch/_5 arc zivon on thu _ibscis_a scale.. a_ntud by points in th_ ._ 1 shows hin_:moment characteristics of a typical control s_urface. In the usual case. 2 shows the _ e_a _" el6v_ltOl_ contr<_l ch. the variations of hinge.. the oi'foctiveness could be predicted quite accurat. _hich i.I_ESIGN By OF CONTROL Thomas A.n of hin_-mem._ l'r. Comth_It would bu rupr_sented by points in t._ard of th.grs in lov.-: the vartat!_i. but the fa_torJ affecting control-surface hinge moments _re not w_ll tulderstood._tic stabilit_ in level l'l[c_ht and is the condition for which ther_ is no variation in stick fore._viding high _aneuverabillty at high speeds with reasonably low control forces.otories..d_. in p_rticular._acteristJcs for _ typical airplane.r0ssed mathe- for _ " _on of th._n_ coefficient _._r_as comb!na_icns of the hin_-_omont 2ar_ters that wouJ._m infowa_tion thut wa_ available in 1941. Over these moderate ranges._ solid line is tiuo bolmclarj for _#.._k force _¢ith sp.moment coefficient with angle of atsack and with deflection are appronimately linear ov_r moderate ra__gos.ho region ab.amut. or never should be ra'.: o:_ci!- b_cc_us. t. alter tile hinge-moment param. _ u. L. botmdar[¢8 plot..altered _ by _he...s.n_ relative _ " in&icates that such an elevator would prow[de u tabl3 stick forces either in level flight or in maneuvers... the regton variation of stick force with nor_l " " "" boiow this botunda-__yreprc_&-_nts conditions for an _msta0!._wic'_.S_ L . The slol_OS of those bo_.?._6_tlv.s 4 co 7._ra r.%bilily.. al to th_ dlst_ll..._. _he boundaries. The location of this po.mi'.].'l_o _.uos of the h!ng_-mu_._r'3_ J}_o..gc-moment p.th spee&..F_ c_. chs...nc. A '.%ra"/ot:_r values for a t_Tpicai plain elevator is sho_m on this to the tw<.3ct/on i_._.. .C-.o (F) . The slopes will varj with the different center-ef--{vrav_ by locations .n h ].D._:._=L:]L_._ va_.e& that a decrease in the nu6_tiv_ valu_ of _C}I/_ is favorabl_ for el(_v'_tor control pro_ided i.e.cL L._s_rablo to decrc_w_.".. since th.and is the condition for zero variation in st_c.q_l.. _'x._2.[sbb.crater in or.J va.a_i Lion to 8_ Ccq.ativJ v<t_lue of 8CI_/_ above the v_luo for a plain rudder wi_'i still .<r..t.._ .o do-'l.o _h..uo of 8Ch/_)5.b _...!...c addit-i.t_.ve values of both of the hin@_-momenb o_.. it s_:>ul_ be nol. The .d.i2__!_r u._ e.p_:._ with ch_Iges in norn_! acceleration _n a pull-up or a turn.tor.b-' _J(.'_-. "'/._r_ _ _. Tho values of thu hin.corom tlio botu_daries tc f the point ropr_sentin#_ the eleve.o.%". the stick forces should be h_ghly stab!o u.I h:owew_r_ is proportior..:' 3 [.e _. :Lg t.k_vo bc_'n ].. in a ¢.rs .?'L._'cr..ticul_r eenber-of-cravity location on a typical airplanu..D el._ount of st. ) of an ("D_3.--ho-_s ow the va!..ated by the methods of r_f_l_nce i f_..l_or<.s_sd much above ze..ining a high n.... g d...± ar_ affected .nat might exist A point tha_ repres_gnts a combinat.D....::rameters can b_-. The ul.:d.L "..._. _':c '_ _. The &ashed line is the boundarb for r/...d bL'r._n.'_'i'J<'l ' :-'-'_"J- )7'0 z'..-.._ ..daries were calc_a]..rlatLon of stick force with normal acceleration.yaucu c of ±ri_f_r... ::.)v_ c.. it maj be desirab!u to..)b_oct is to move th_ point f_.._]!?osi._u]_i-_].ih¢_t ..r._iL'.ion of hingemoment-p.. a of -.-uvering stabi!i.om!.l.Otc]"[3[ . d. Most balances cauou a decrease in the neg_tJ.1_uct tJ!3 c_._:d_i.T.:. a oc. it also seems d...rol d.rc_ w-.nL:ro! ctLrl'_:._3 ._' 0Ch/ch_ // for t_.haL it _s not accomp_inod by too _au a decrease in the negative v'&lbe of _Ch/O5 .-' " _ " ba]..c.. nts conditions for a sfl_ble acuel_ra_c-I _' and.rti_er from the bound_rl .bivo v_luc8 of l_Lions w_ th c(.on _..D]._.. :"_o nc6._ forc....:¢i... O_'":_ . t_ercforo.d.r_-?.:: 0 .i-...L.u by t. OX i. _ above this boundary represe._n '_' h .ler to move the point reprssenting the elevator farther from. ._balance .....Ol." ! .. 'v.TL (A) _i.L.e ru&d. +..:. . In the o_se o±' the rudder._._tors of th. kinds . For l_g_t aircraft._..:.m_.'_C_ ..._:_otors._i.. addition of aerodjnamic balanc_.3J a _).. mainte. .. /..nd...l_free Lhis par_letur . Th_ w]L._:_..z6 region b_!o_¢ this bcund&ry would g_ve an unstable variation of stick fc.. ! ]:.. _)_ the sD_mm:_._of the required ftunctions of the control surfaco but. horn ba!_nce ant! the i.. pb/2V = 0.an aerod$'n_lic chord to 40 percent of' the _ean _erod_'n_mic chord.ilcrun chord _ as a fr_ction of 5he wing chord is p]._l_. and the Frlse balanco probably are of . and the sealed interns! ba!anco (E) have intermediate effects. the force per g varies from a high positiv_ v'_luo to a _u_1._s_d by a consts_nt increm<_nt throu_out tho con_r-cf-gr_vi<. The shielded._m_al acceleration .y r_ng_' bj" attachin£ a bobwelght to the control stick.[_ c.o_t_d against the ail_ro_ span as a fraction of the wing s_mispan. The force __.of-gravity range.3 c_nt_gr-o£-_ravity range.lat should be capabl_ l oi' meeting the rocuirem<:n_./_5. it _a2 be dusir[.ri. The most desirable t2-po of balance for a given control stu_f_co can be sel_ctod only a/'ter _. while the addition of an tulshie!ded horn balanc_ raises the forces tbrough tho 6_enter part of the center._t suited _o o__..gati_ vaiu_ as th_" center of Craviby is mo_.%l light alrple_no _:_ffects the s_!ck force p_r tuni_ no. For th_ plain _levator.ic bala_io_ or a bobwo iGht.'ise balance c_nnot be ropresente_ very satisfactorily by this type of chart because their characteristics are very nonlinear.!_forces low throughout th._ors or rud.vators. In g_neral.1 n_.07 with no regard to control force.a_ic_d . 0_h._ elevaCor of a typic. Their effects on control forces however are similar to the effects of the s_-mmetrical From overhang balances.0valor.bl. If _-% lift l'l_>q_ is to be used inboard of the ailerons. Fi._. such as the beveled tr_ilin_ edge (. Th_ problem of aileron balance _ s influenced to a large _:xt_nt by the din_nsions of the ailerons and by _he deflection r_6o. ra:.commonly caile_ the stick force per _ over a range of cenber-of-cravlty locmtlons. of ease of Funufacturo and maintenance the considerations tu_shlelded born balance (A). for !.ne_'g c::_nb<._._akos t].lcrs and th_ _k'is-_ or syn_v_trlcal overhang balances are _°. oes_ .raft._. Figure 5 shows con_binations of SpoA%. but simil.evices.ca! overh_ng balances (C an_ D).b. the s[_:u_trlcal overhang balances (C alld D). chor_.u suited to ailc_-ons.sr _.akin_ a thoroug_ stud_ .greato_ interest in light-aircraft design... the unsL__e!ded horn b_!ance is b_.27 causes a negligible change in _Ch/_k_ for a _iven change in e)C). az_! deflec_Jon r:lngo of plain ailerons t. . is ev_dt_nz from th_se resl_its _hat a considerable amount of control over 51_e force p<_r g is avnilable through th_ use of nerodyn_'_.%_ effects would be obtaine_l through the u_:o of _ bebw_:: ght in oonJlunctien with either of the ba!. The e_'_'_ct of a bob-' weig_ht is lllustra_ in figure 4 for the case of a plain 6:].i from 20 percen_ o±' th_ _. The addition of a rotund overhang balence _.ght air. it. to resbrict ( .T_re 4 shows how the addition of aerodynamic balance _o th. but a much lower force is desirablu provided t_e friction is not so great as to prevent the ailerons from returning to their tr. that the aileron span is to bc_ shortened to .07 for a typical airplane with no aerodynamic balance on the ailerons. wide-chord aileron._loased. The dashed curves show the speeds at which a stick force of lO potu_ds is required to give a pb/2V of 0. principally because they exhibit undesirable floating tendencies in sideslip. Figure 5 indicates that the requirement pb/2V = 0. long-span ailerons. however. the stick forces at a given speed may vary considerably for different aileron configurations capable of providing the same value of pb/2V._quiremcnt can b_ met at 150 miles per hot_ with ailerons having a chord of 14 p_rcent of the. If a lift flap is to be emp]. th.07 ca__ be met _re considerably lower for the wide-chord. Although it is possible to meet the requirement of a pb/2V of 0. Assumo_ for cx_. Careful design of the ailerons with regard to their hinge-moment characteristics may. In order to meet this requirement at a speed higher than that given by the dashed curves some aerodynamic balance must be used. Figure 6 is presented in a form similar to that of the preceding figure. for a given stlck-force requirement.07 can be met with a very small aileron spa_ provided the aileron chord is of the order of 0. (reference 8) has not shown much promise for flight experience such ailerons. It sho_d be noted of a force of' lO pounds that the speeds at which tl]o requirement for a pb/2V of 0.oy_d on the airplun(_. short-span ailerons than for the narrow-chord.28 the ailerons to a small span. if there is no obJ<_ci_ion to the use of an aileron spa_ of 70 pu_'cent of the wing semlspan.%m position if they are deflected and then r_. or._ aileron span probably would have to be shortened considerably. w:_ng chord a_d a total d_flection range of 20 °. there would b_ no ne_d for aerodynamic balance with such _ _il_ror_ c_r_figuration.30 to 0. The wing spaz_ of the airpla_e is 35 feet and th_ aspect ratio is 8. For most li_t ai_]_lanos._mple. Although the pb/2V requirement can C be met with a small-span.07 with widely different aileron configurations. the spewed at which that requirement can be met depends largely on the aiL:_ron configuration. in the total deflection range to above 40 ° are of very little benefit because aileron effectiveness drops off rapidly at the higher aileron deflections.40 of the wing chord and the total deflection range is about 40 °. This is illustrated in figure 6. A forc_ as high a_ 30 pounds may be acceptable. conversely. Increas_. and the solid curves which represent total aileron deflection of 20 ° and 40 ° were taken from the preceding figure. The stick-fol_ce value of l0 pounds is rather arbitrary. provide some improvement in this respect. the r. For uxam_le. ) A collection (r_f. the aileron chord would have to be incre_aso& to about 37 percent of the wing chord in order be obtain the desired aileron span. 'l_e slot-lip ail_r(_n a'_ shown is not well adapted for use with flaps retracted. Incl-eascs in both. This will necessitate an increase in the deflection range. they w_r_ usud only with the flaps deflected.:_renc_ ll) of slot-lip ailerons on an airplane. In one full-scaLe installation (ref. The use of the slot seems to overcome th_ objectionable characteristics of th_ hh_god plato and th.'d abovu the _o'_ng surface. Th_ plug aileron (references 15 to 18). when used with a slotted flap. This device is designed in such a manner thab a slot through the wing is opened as the plug is proJect_._wardmowm_.r and th._'._ded if the requirement.rthis confi&_aration. Spoilertype devices seem best suited to . unless the devices are located very far back on the wing._. seems to have satisfactory characteristics with flap_. and._rence 19) has been made of the me Jr pertinent !__formation _ven in th_ _-_rious reports. Aerodynamic balance would be ne._flection range would have to be increased to abcut 40 °. is to be met at a higher speed.devices. . Figure 7 shows four such.the hinged-plate spoil_. the d._ _i[_h_r down or retracted.._-arc spoiler .because the position of the retracted flap prohibits anj d<)'. For this configuration.t of _h_ ail'..s that: now are being released.. the aileron chord. or possibly.are considered to have trades!table chsa_acteristlcs in that sm_ll projections of these devices abov_ the win_q stuff ace may'produce little or no rolling moment while slichtly larger projections majw cause very large rolling moments. a stick force of lO pounds for a pb/2V of 0. It is evident from the foregoing example that aerod2namic balance may be needed if short-span ailerons are used. If the d_flection range is n_intaine& at 20 °._ re tractable-arc spoilers. F(_.07 the r_qulrsment could be met up of a fo±_ce of to a sp_ed of l0 potunds for a about lO0 mJles per hour. retr_. h.vthat operation._ron and therefore a rather complicated linka_e would b_ requlred for its operation._9 35 percent of the w:_ng semispan.)wov_.r. Data on these and other devices suitable for use with full-span flaps are contained in repor_. Lateral-control devices that can be used in conjunction with full-spaa_ lift flaps h_ve been of inte±_est fol_ _uy _ars.07 would occur at about 65 miles per ho_. In these respects _h_ slot-lip aileron is satisfactory with flaps down. If the aileron chord is maintained constan!-. a noti('oable tLme lag _ay occur between operation of t_ devic_ and the response caused b.. (See references 9 to 18. pb/2V of 0.ctabl. The first two el' the devices show_ in figure 7 .such arrangements. Conventional ailoron_ on ch_ trailing edges of the flaps were used with the flaps rctracix_d. radians pb/2V P b V 5u 6cl helix an@!e developed rolling wing velocity. or flat plates prc_jected into bho air sLrcu_. radians per secon_ span. feet span of control b S surface at h_[nge llne. a spoiler. root-moan-square line. SYMBOLS variation of hinge-moment attack coefficient wit_i angle of 8Ch/ 5 Ch variation of hinge-moment deflection hinge-moment coefficient degrees coefficient with control (H/qcs2bs) angle of attack. degrees decrees airspeed. but it is felt that the available information on lift flaps. H q C S c _ntrol. dynamic foot-pounds de gre_ s pressure. co behind hinge. fo_t in roll. spoilers. Very little work has been dcz_e so far on the development of devices specifically for glld_ control._t chord of c_)ntrol surfo. . Such a device may take the form of a simplc lift fl_io. hinge moment. and v_rious kinds of aero_lyrzunicbrakes (reference 20) can be utillzcd _n the design of glide-control devices. a device capable of controlling th3 glide path may beccme important._ to add parasite drag. 5. for l'ght aircraft. an_ the ailero1_. the rudder. feet foot per second deflection. pounds per square f.surface deflection. upward aileron @e_nward aileron deflectlcn.3o LT"I In addition to the elevator. 'neModel Equipped with a Rudder HaviI_ Positive F]. 3L07. (Prospective TN.i_ation of 'bl_ Lateral _ Oscillati<. Harry.cl System. Research Denartmont of Lateral-Control 5. M. Melvin Flight Investigation of the Lateral Control Characteristics of Short Wide. 1943. Weick. _'_CAAC_ _o. #. F2A-2 Airplar_i_ with Full-Span and Slot-Lip Ailerons. and Sterr_fleld._.1 Surfaces. John G.: _'i. Lowry. Greenberg.n _ Wing Full-Sp_n Fl_p Consisting of an Inb. 6. J_'seph W. F.[le_}n Test Data. Ail_roi_s and Various Spoilers w:ith Diff_n'ent Amounts ! of Wing Dihedral. NACA ACR No.ns ._I' an Airplane with Free Rudder ww__]hSpecial Reference t the Effect of Frlcti:)n. Richard H. Greenberg.. -Fred E.. F. Rog_!!o.RF_CES 1._tiug Tendencies _uld Various Amounts of Friction. 7. . Rogallo. 1944. 10. 1943. and Spans.ght Tests of Slotted Flapc _uad Tr_dl!ng-Edge ARR No. : Wind-Tumlel Inve_tig_t_on_ cf a Plain and a Slot-Lip Ailer. NACA ARE. and Sternfield.. and Schuldenfrei. NACA ACR. T_. .) (Compiled by Research.: Collection of Balanced.-A._rd F. NACA P_ N. Marvin: Wind-Tunnel gatlon of a Plain and ri Slot-Lip Aileron _.. April 1941..r Unshleldod Surfaces. 3L08. March 1943. 9. Bartholomew S. _Alli 194!_. and C<. NACA ARE No. and Sawyer.n _n_ a W_ng a Full-Span Slotted Flap. Francis M. 1934. NACA with InvostiwitA a an ii.II): (Prospective TN. NACA ARR.. Rogallo. " • N. Wind-Tur/_uel Data on the Aerodynamic C. Harley A. Wetmors._ry ReSume Control of Hinge-Moment Data f:._wler :_d Outboard Slctted Flap. Maggin. 1943. No. ! Horn-Balanced Thomas A._ntr. Richar_ I. :. Sears. 2. _5oule.ugh. NACA Rap. : A 8. I 4._pl_.) Sum_. Harry. 4BO1. Leonard: A Theor_tlcal Invesalgation of Longitudinal Stabil:ity '_f Airpl_mes with Frse Controls Including Effect of Friction in C. Bernard: Ex_erlm_ntal Veriflcat:i<:n o£ the Rudder-Free Stability Theory for an Ai. 3F19. Leonard: A Theoretical InVest. June 194!. M.haracteristics [of Airplane Centre. 494. . and Lowry. Deflector. Lowry. F. July 1942. 706. F.. Wind-Tunnel Investigati_Jn Aileron on a Tapered Wing ARR. (i.: of a Plain Aileron saxl a Balanced with Full-Span Duplex Flaps. June 1942. M. : A Study of the Application of Data on Various Types . NACA ARR. Fischel_ Da+_ in preparation. John G. : Wind-Tunnel of a Spoiler-Slot Aileron on an NACA 230]. No. Robert S. NACA ARR. NACA ACE. NACA ARR No. John G. Francis M. : Wind-Tunnel Investigation of Spoiler.2 a Full-Span Fowler Flap. N&CA 15. Rogallo. Investigation Alrfoil with and Spano. 13. Robert B.. and Ivey. M. and Purser. 18. Rogallo. July 1942. Bartholomew S. Harris. 1941. Flap. Carl J. Jack: Wind-Tunnel Investigation of a Full-Span Retractable Flap in Combination with Full-Span Plain and Internally BalanCed Ailerons on a Tapered Wing. F. Dec. Lowry. 1941. Jack.. : Wind-Tunnel Investigation of a Tapered Wing with a Plug-Type Spoiler-Slot Aileron and Full-Span Slotted Flaps. Francis M.. Wenzlnger. 1941. Paul E.: Collection Full-Span Flaps. and Rogallo. and Slot LateralControl Devices on Wings with _h_ll-Span Split and Slotted Flaps. 17.. Margaret for Lateral Control with ) F. 1943. NACA ARR.. and Fischel.of Flap to the Design of Figi. M.32 12. March 1941. . Rogallo_. :_f Test (Paper 19.. 3H23. John G. NACA Rep. Purser._ "" 20.ter Brak_s. Rogallo. 16. Thomas A.. and Swanson. a_id Liddell. No_. NACA ACR. : Wind-Tunnel Investigation of Plain Ailerons for a Wing with a Full-Span Flap Consisting of an Inboard Fowler and an Outboard Retractable Split 14. Paul E. : Wind-Tunnel Development of a Plug-%_jpe Spoiler-Slot Aileron for a Wing with a Full-Span Slotted Flap and a Discussion of Its Application. t_) stability .Hinge-moment characteristics DEFLECTION.008-. Ch © E_ .2 .HINGE-MOMENT .004- PLAIN _O04SRT%BILNE _//_/_BOUNDARY FOR LEVEL FLIGHT FOR MANEUVERING " 008 _// _012--.I 0 ANGLE -. DEG -5 0 5 </i -.Relation of two hinge-moment boundaries for elevator parameters control.3.I OF ATTACK (7. of a typical surface.3 COEFFICIENT. .004 (_Ch 0 .2 -.008- .4 I I I I I I I I I -25 -20 -15 --I0 -5 0 5 I0 15 _. 20 DEC control 25 CONTROL-SURFACE Figure I.3 15 -.020 /// I _/_BOUNDARY UNSTABLE REGION I I I I _016 _012 -.004 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Figure 2. - Conlrol-balance arrangements moment and parameters their effecls on hinge- 16- PLAIN ELEVATOR AND BOBWEIGHT ..G..004 BALANCING TAB . -008 .SHIELDED HORN BALANCE FRISE BALANCE i--I o E_ @ ROUND UNSHIELDED HORN OVERHANG 0081 . -004 bCh/b8 0 .012 ® . 70 MAC balance maneuvers and Figure stick 4..OOB -i " -.... LOCATION.004 _Ch/_ 0 BEVELED f {j " "": ®_ _) _EALE_ ..- Effects of elevator forces in accelerated of a bobweigh[ on elevator for a typical light _ir_)lane.NTE_NAL BALANOE TRAILING EDGE -. 12PLAIN STICK FORCE. ... i 20 25 50 C. POUNDS PER g a ELEVATOR_ " NATtONAL ADVlSOnY COMHITTEE FOR AERONAUTICS m HORNAL B ANCE--- \ "J ] 55 -i 40 0 ROUND-OVERHANG BALANCE t 15 r F ..004 _ i -. SHARP OVERHANG PLAIN CONTROL SURFACE Figure 3. 07.4 I .5 ONE .AILERON WING o CHORD CHORD MAXIMUM ./'75 _ ffO0_ _ (MPH) WING ASPECT 20 ° FORCE _ pb 2v SPANRATIO= TOTAL --I0 . _) I .._us speeds a value of pb/ZV equal to 0.I ._ _.07 55 8 kBS FT ! o2-] i I_1_ F LE-GTION .4- TOTAL 20 ° DEFLECTION = _u +_d 50 ° 40 ° 30" .0 ADVISORY FOIl AEgONAUTICS AILERON WING Figure 5.Aileron SEMISPAN CONNITTEE configurations capable pb/2V equal to 0..6 I I .50 I .r{:oo_ 10 _)_um.Configurations of plain ailerons capable .7 .9 I .2 .I 40 ° TOTAL DEFLECTION-'-" I I . .8 .5 _ _ i 1..0? with a stick t: .9 NATIONAL 1.II2..7 I .>t7providinc. :_twlri.3 SPAN .8 ._ .2- (% .IO 0 .1- I I I I I I I I I I 0 .. of providing a value of AILERON WING CHORD CHORD SPEED It \i 50 \ \ /.6 .Js for a typical light airplane. .I SPAN OF WING ONE AILERON SEMISPAN Figure 6..4 OF . 0.5 .5- .. . © " I t (a) HINGED PLATE SPOILER (b) RETRACTABLE-ARC SPOILER f C % (c) SLOT-LIP Fi_ure 7.- AILERON lateral-control full-span devices flaps. (d) PLUG suitable AILERON for use with Spoiler-type NATIONAL COMNITTEE FOg ADVISORY AERONAUTICS . iJ. C It is the ptu_pose of _.!.s keenly felt It _.Cluc_. 4 Comparisons of hin_o-momez:t parame-ter..'. _._diction control of forces _nd to show he'# some o'_' the re':ulto may be applied to the problem of s_!ck-.-SI_FAOE By Charles W.: data +h_t .01.ec Laboratory._ent cD'u-_'sc'.._cs of flmi%-.this methnd _ith e_rperiment. aspect ratio. is based on r'e.rch tc d.:. primarily a fs_nction of aspect ratio.al tails of aspect ratio great(.g-!.. especially _ns_far ss the prediction of varlaticn of control hinge moment with tall angle of attack was .m_.onthl_ problem bein_ done at the . slm.-:ve]op a method of prodlcti_L_ c.'._ccnc.i_c the'_ry.e_s in t stability due to freeing the controls.free stability for " _* a._._.t some ch_.rc._its for aspect ratios ]e.d 2. _ var_'_tioi_s in. c:ntrol--sur_'ace chords._ predicted by.Jass..-aues pr<.his.r th'._ON OF CO_FSR.._el_t res._ers "_.ry of the research the NACA is ccnductin_ on the p'.racy of prediction _2.'_.nt character!st. Frick. aspect ratio from soction data has _ee_1 under study in the laboratories of the NASA for a n_mCoer _f years. _...t is I_o_ <s "lift:[r.racy for prnlimina_'y des._<ht use.o _..'ce as the aspect ratio decreased.ng this vortex iatt..uCo in -k.ectod._:. Yt is the basic p_rpose cf this rose.r.cn )_._(. " .. a_..y the 25-percept chord line of the su._ -<b.e..'_o=m. o h chordwi_c dis÷.cc._i'acc._'s ago the NAC.._. Jr.0. yea.._.c:._!ts of here at the Langley Laboratory of The prob_-.._l res_. It was fo_d that the crror Was of appreciable magnitud.ct_on was caused by the fact th_.ntratod in a single llne._..peper fie present a <ener¢3. sho_.. tailor ratio.m of control suz.__n especially for horizon+. s.aft The research and also ans!ysis that is presented ._ntre]._i.]..he scot.<:_ S..--surfacc h.__f .wore fecund to be of sctlsfactory scot.'e_'_r. __th trailing vortices bein_z shed as th_!_ tip is approached._J_.j I_A_DiCTT. nces 1.erlst!cs f_om ba.dictiT'_g the hlrg_}-mcF. usua..ublishod a mst}-od of hin(<e._nc_ this paremeter ___eatly !n.s he ].. ho_.u_niced th.mom_:nt predict__cn _hich _as based on ._ V... the n_ed for improvin_ _.._vc_a]..the t error in pre.(.) In this thoorly the vortices which represent the lifting capacity of the tail surface s._s being ne&'}. This is sho_cn in figure i by the vertex pattern on the !eft. sections which aircraft desi_.._c:._d some error !n the predictions which bccar. .rlbutlon of lift._.md airfoil .:i!! sccovx_t for the _ig. Predicti_n of hinge-moment charactc_istics us_. ..d..._as'. patt<_._r the hinge of th._ of attack !n going from infinite aspocb ratio to as_._ct....e sgr_-_oment._flcctJon shown on figure is deflected two tjpcs of chord.?n that the iift-_n_.i._nce 3)_ a further m_alysls of the problem _.ing-lJno and 11.L. (S._• 2_ 4_ . Comwaring the cross-hatched areas for l_fting-l._:.mg-ourfac_ theo!_y._oct ratio 3.dd_.[c__n the exp¢_r!menta]. shows a comperison cf the chcmge in the r. theft is...in:.on._sh as dcte..ng!e of attack from lifting-m_rfacc theory is given by the cqu_. the lift distribution along the chord of the wing is altered with a resulting change in hinge_moment-producing load over the control s_rface._-_'._ a c_.oq_Hn_ h_c _._to deflection of th<_ control which has a p_>_k ov._l .34 With the aid of a new application of the vortex theory of wings.zasmade. CcmDuri_onu b_..) . the induced downwash is calculated by th._._. _e 05ii. chanc e _n asg.. rat_o._.v provi!_s a correction to the chordwi.fercncc 4.on of control hinge moment w__th tail _. it amounts to 60 p_rc_nt. _._ loading due tn.in proportion to the local lift. The c. the lifting surfae_. _ • ._dlng du._.llftin. loading resulting fro_.IjD_c of the voytc.nS tile airfoil add_tJ. hinge moz_mt.s that which would.e chango with aspect ratio in the control d.xpression for Ch_ . this analysis uses the vortex pattern sho%"n on the right of figure 1 wherein the vortices representing the lifting capacity of the surfaces are distr_ibutcd ov:_r the tail .in th.. With this vortex pattc_rn. theoi_y s_m'_._"_ t re._t_. the variat".urface thsory.._. it c'_n L<_ seen that the. result a Jf th_ [ncr._go in aIV.: control.z-s_Irfacc correction to the chordwis_ loading is n:_.6_ --_.Incl"casi_g th':.r du. called lifting-s._ho_dW_.._6 r:._en the control is to _. and lii'ti_-suz_ace theori_'s.-.0 for . If the down_. developed by Miss Doris Cohen of the Langley L_Jboratory (refer. .Ana.se loading _iven by !_Ifting-l_ne thoor?/ as _hown by th_ lower crosshatched u_re. Figm_e ?. below the diagrsm on the right of figure 2._:mcthod ¢_ Miss Cohen at a number of points on the tall by summing up th_ contrlbu_ion of all vortices._nnined varies along the chord of the wing an_ the streamlines do not conform to the profile.rn for this type of cho_dwise ].. chcrdwis_ lo'. th__.%f t..ef]_/re o on th_ right of fi_ur_ 2 Knowing the change in _._ly thu s__n.. a.: 3. _.__. an{-_le of att_%ck.::t_sDect rnt_ o of 3. Essentially..?. surface. The change is indicat<_d by th_ cressh%tchcd area. At an aspect ratio of O. of COLli'6_b_ compute the chang_ iD. _ .tion cho_._.l.vu!tin_: "" " I "_' tire ch_l_e ._-_heory correction _uuo_nrt_ to only 20 p6rc_nt of the total ch_n&_e in the variation cf control hinge moment with aisle of attack while ._ Lift in duo to control deflection were obtained by simply .basic control lift associated with the change _n aiiTo!l camb_.llft. values data of show Oh~ determined by this equation and rem_'kab].Aing are '?.wise loa._ho_.. the effect is more _owerful. This is shown by figure h. say.075 or a rearwa_ shift of the neutral point of 7. that this method of controlsurfac_ hinge-moment prediction gives the d_signer a means of determining the basic geometric characteristics of the horizontal . more rapidly than Chs.0 to 3. a re_N_xl shift of the neutral point of B percent of the mean aerodyn_ic chord. The hinge-moment parameters shown are those for a 30-percent chord elevator on a tall plane of NACA 0009 airfoil section and elliptical planform. c1_mging the _sl_ect ratio of the hoi-izontal tail from 5. say. 5000 pounds. that is. You will note that the d_fference between llfting-lln_ snd liftlng-surface theories become of significance at low aspect ratios. this means that the equation shown on figure 3 e_ressing the relationship between hinge-mument characteristics for aspect ratio _. where control balance is required to reduce maneuvering and landing stick _orces. Ch_ approaches a value of zero at aspect ra_io 2. Figure 5 shows an example of how it may be a_plied to the estimation of stick-free stability° The resu!t_ sho_n have been calculated for constant stick-flxed stability. therefore. profile characteristics. become less negative. the change _n chordwi_e loading due to aspect-ratio change is of the s_me form for lift produced by control deflection as it is for lift Obtained by change in _ng!e of attack._ percent of the mean aerodynamic chord. This is shown by the cttrve labeled 3_-_e_ent-chord elevator balance.0 to 3.5. 2_00 pounds (_ss _i_t. It is evident.0 results in an incre_Ise in stick-free stability of 0.35 _adl _ngXe of attack. This amounts to 20 percent of t_e stlck-fixed stability of the assumed airplane. /5 ¸ which is usually of very small m_nitudc. The curve for the unbalanced elQvator may be considered to be that for a personal aircraft of. _n other words. changing the aspect ratio of the tail frc_ _. This new method of control-surface hinge-moment prediction is of very great use to designers. As you can _ee. For a larger aircraft of.0 results in a gain in stick-free stability of 0. From this equation it may Be seen that the aspect-ratlo correction %o Ch8 is always less than the aspect-ratio correction to C_ so that as the tall aspect ratio decreases we may expect Ch_ to iucreasep that is. that is.03. Ch8 is affected to a smaller extent. Fortunately. and for any other aspect ratio has the same form for both l_fting line and liftlng-surface theory except for an additional term (dChS)LS . As the aspect ratio of the tall decreases. This amounts to one-half of the stick-fixed stability of this airplane. For this airplane. Ch8 is affected to a smaller extent. bin@e-moment preaictlon is This new method of control-surface of very great use to designers. c1_mging the _s_ect ratio of the hoi. profile characteristics. 2500 pounds (_ss _i_ht.5. and for ar4V other aspect ratio has the same form for both lifting line and lifting-surface theory except for an additional term (&Chs)y_ . that this method of controlsurface hinge-moment prediction gives the designer a means of dote_mining the basic geometric characteristics of the horizontal .0 to 3. The hlng_-moment parameters shown are those for a 30-percent chord elevator on a tall plane of NACA 0009 airfoil section and elliptical planform. that is. say. For a larger aircraft of. say. This is shown by the c_ve labeled 35-percent-chord elevator balance. You will note that the dlfference between lifting-line and lifting-surface theories become of significance at low a_ect ratios. This is sho_n by figure h. As the aspect ratio of the tail decroases_ Ch_ approaches a value of zero at aspect rab:io 2.0 to 3.v expect C_ to iucrease. more rapidly than Chs. a rear_Lrd shift of the neutral point of 3 percent of the mean aerodynamic chord. This amounts to one-half of the stick-fixed stability of this airplane. that is. For this airplane.35 _adl angle of a_ack. which is usually of very small mJ_zitudo. Fortm_ately. this means that the equation shown on figure 3 expressing the relationship bet_en hinge-mument characteristics for aspect ratio _. The resttlts sho_ have been calculated for constant stick-flxed stability.0 results in a gain in stick-free stability of 0. It is evident. the change _n chordwi_e loading due to aspect-ratio change is of the same form for lift produced by control def2ection as it is for lift obtained by change in mag!e of attack. changing the aspect ratio of the tail fram 5. therefore. As you can see.5 percent of the mean aerodynsmlc chord. From this equation correction %0 Ch8 it may Be seen that the aspect-ratio correction is always less than the aspect-ratio to C_ so that as the tail aspect ratio decreases we ms.03.0 results in an increases in atlck-free stability of 0. in other words. This amounts to 20 percent of the stick-fixed stability of the assumed airplane. the effect is more _owerful. Figure 9 shows an examplo of how it my be aoplied to the estimation of stick-free stability.075 or a rearwa_ shift of the neutral point of 7. 5000 pounds. where control balance is requil'ed to reduce maneuvering and landing stick _rces. become less negative. The curve for the unbalanced elevator may be considered to be that for a personal aircraft of.izontal tall from 5. that is. C_ _gle attack.35 tail necesso_y _d.s c r ip t: L__ i! ft_!. square of wing./ of control hin.dth control deflection pitching-moment lift iDrmal CL AR M q C • Cm CL coefficient (L/qs) per uuit lift (M/qcs) coefficient pressure ratio moment..n.control g_. S L V airspeed..ents. density of of per second. for ti_e attc.Ee-moment coefficient . pressure.ad!.e . .:'. force control-free stability desir_b].. K ceustant S:ub.:'.. c_ic foot slu{_s per degrees C. pounds foot air. SYlV_OY. foot-pounds pounds chord feet per square feet foot aerodynawLic area.S variation Ch_ angle variation Ch 8 of control of attack hinge-moment coefficient _th tail /-L.. su_f._c. coeffioient aspect pitching dynamic mean _ng lift.!nment of desirable. No. 1942. 721. 1941. 5B05.oad Distribution over Airfoil Section w_th Plain. H. NACA_¢ No. . 2. _.es._ilton B. . Julian: Calculation of the Chordwise I.37 REFERENCES !. 855. or _orially Hinged Trailing-Edge Flaps NACA Rep. S_lit. Jr. and Sears. 1938. Allen. NACACBNo. Crane. Cohen. 1945. _. Eobert M. • Dete_.: Computation of H_r_e-Moment Characteristics of Horizontal Tails from Section Data. Doris: A Method fer Determining the Camber and Twist of a Surface to Support a Given Distribution of Lift. Richa1_l I.. No_ 63_.. NACA Rep.ination of Control-Surface Characteristics from NACA Plaln-Flap e_udTab Data. SURFACE LIFTING -.LINE Horizontal-tail C vortex patterns.oo .Correction to the variation of elevator hinge-moment with tail angle of attack due to change in aspect ratio.AR-_O P CL _P GL AR.LIFTING LINE LIFTING SURFACE C. NATIONAL CONNITTEE ADVISOI_Y FOg AERONAUTICS . Figure LIFTING I..AR'3 i r X i GL= Ch=" Ch(¢--_)* (&ChJLS Figure 2. . _ I--_ _ ...0060' _ ."_ DEFLEGTION..... parameters with . 1 .... LIFTING LINE CORRECTION It • I_ \ I_\ I_ \ r UFT ADDITIONAL DUE TO _'... t _ 1 t NACA 0009 AIRFOIL WITH 30_ CHORD UNBALANCED ELEVATOR Figure 4..........- Variation of elevator hinge-moment h(irizonLal-tail aspect ratio....- Correction to the variation of elevator hinge-moment with elevator deflection due to changein TAIL ASPECT aspect RATIO ratio.. _ .... I .:5 "I . EQUAL TO SUM OF BASIC CAMBER CHANGE LIFT AND LIFT DUE TO EFFECTIVE ANGLE OF ATTACK CHANGE P CL BASIC LIFT FROM EFFECTIVE ANGLE OF ATTACK CHANGE DUE TO CONTROL DEFLECTION..r-. ..Ch= ) ÷ ( ACh6)LS _z_Lu-e 3.. _ r LINE i i .-_ ...... I\ _ _ -ff _ Jl "__ CHANGETO DUE _....-t EQUIVALENT _'....°_'...._..-. e_ "n6 _ ...... HORIZONTAL o ---_. NATIONAL COMHITTEE AR 3 ADVISORY FOIl AERONAUTICS Ch 6" Chs+ =¢s ( Ch=.. r z -0060 _/-_ .(")l_)n (_ ---/ _11 ' ' fl : ....kl b_ . AR. UFTING SURFACE _\_-_ ::_ CORRECTION ! ANGLE OF ATTACK CHANGE.'/-'...'/" G:///. LIFTING _ -oloo _ o-..-UFT.' ! ___ f -_2o • -0040 o I.. !:. _-......_ ch...........NGNE __! U i --q _-.. 10 A(_-r--). FOR AERONAUTICS .02 /./ V"hs / _ .04 / I FNO =L_Z _z I. NACA 0009 AIRFOIL WITH 30% ELEVATOR / _ u.06 .--t ._l.z tu .i Figure o / 5..= _o . 6 Varlation wlth aspect ratio of the decrement in stabillty due to free controls. / /.Oe °osTo_ _ "/" .. O_L 9 dC • K .- / 3 I I 4 ASPEGT 5 RATIO ! t i .12 oo . NATIONAL ADVISOI_Y COMHITTEF._---. .38 (i -..'i : . could not be _ta!led iu t_u_s.orated in a light airplane to increase its s_fety. po_sesse_.39 A _._ By Paul A. Fi_u-o 1 shows the airplane _s it appea_-ed originally together _.ady in service.!ng the nose. As _s mentioned before. _ _hile net stallproof in straight fli@ht. The stick force characteristics of the _odified airplane likewise sho_zed a reduction in the effect of powe_= on the _tick forces. The method originally chosen to _._Ith the mo&ificat!on_ _.L:ationwas an ai_!ane of which was spinproof. both with power i era _md power off. a definite step in this _irection w_s t_=en. Stalls from straight and level flight __th the original airpl_ue resulted in rapidly. The stalling characteriotics of the modifle_[ airl)lane were considered superior to those of the original o. Elevator angles above which instability occurred arc shown by the cu_'cs of fi{_ure 3. The re&uction of the effect of po_mr is also apparent. Note tho difference betwcon the newer-on _nd pow_'-o.ud net basicslly to alter the appearance of the ai_lane. the effect of power . While this _s no: completely achieved. of The region in which the lateral instability occurz-ed for the original airplsne and in which the longit_ine_l instability occurred for the modified airplane was defined by the u_-elev_tor position. consisted merely of a m_l& d_._rplane_nd _e for the mcdlfio_l airplane. rplaue in that the stall.__ conditions. eupe::ior _tallJnG characteristics as compar_. Fig.ng 19'$0 and 19&'! a series of modifications were inco_. to s_e value below that required to stall.t this investi._. T_le final -.vC_ INV_LTION TO INCREASE SAFE_ OF A LIGET _I_L. 19 ° for the origins! a. them simple enough to be _ readily incoz_orated in airplanes alr.te the effect of powez° (_nthe elevator angle required te stall _d !inZt the up-elevator trove?. A few of thcs(_ ch_m_cs _ im_ediatoly apparent but oil the changes _lll be d_scribo& in detail later. diverging lateral oscillations which could not be controlled by the ailerons.Z.-opp. H_ter _ur_.d with the original alrplsn._re 2 shows the stick-fixed static longitudinal stability of the original and modified air'planes. It was intended to hel(_ the changes to a minimum._ke the airplane sial!proof was to elimin_. to make.'hich N_ero ma_e tc it.2es0_. The increased stability of the modified a_rpl_ne is apparent despite the adverse effect of its more reazN_ard center-of-gravity pocition. e. h'_. The modiflod airplane with the li_dt_d rudder could be taxied s_ti_factozily with th_ use of b r. The violence c.s 2_._ for a contur of _ra_y of 30 percent of the mean .uce of the wing was chaBged _'rom 2 ° to -1.equlred to make the modified a_rplane stall-proof in strai#_t flight because of the up-e]cva+or d_floction required to make a _-point landing. It was not practicable to limit the up-elevator deflection to the exte_t _.rns with either u power on or power off._vators out of the pro_ell'_r sli2stream all tended co increase ta_ elevator angle roqulred to stall with power on.th po_rer on or power off.2o _and t_ washout was increased from 3_tc _ • The thrust axi:_ _.vel to _15o co_ld be held in a steady spin to th_ left.ucrease&.l v.rs we_e move& o_t of the oropeller slipstream _nd the ar_as _nd aopect ratios of the horizontal and vertical t_lls _¢ere i. T_G original a_rplano could be stalled out of a tu_ with power on or off..erodynam!c _:hor_l _d wa:] reauc_a _u as t_e cen_r of _ravity move& back 7 percent. The eliminated as planned but a large reduction center-of-gravity perdition for a given loading not of the airplane was moved back consider_lybccause of the increased weight Of the tail but the range of lo_dlngs was affected.r_ilod. rec[ucing the incidence_ and m_ving the o].s n_d no loss . Increasing the wing washout eliminated the lateral instability at the shall._._. The Incide._v_tor deflection required to Droduc_ pitchin_i v_locity in thu turn.4O was was not completely effected._r acc._s _ep_ed 7 ° a_d the rudder travel was limited to +15 ° • 'E_e _levatc. Increasing the aspect ratio of the horizontal _d vertical tails improved the effectiveness of the surfaces whil_-_ the add_d ar_a increased the static longitudinal stability _d the (iirectional stability of the airplo_e. If th_ left rudder was d=creased to 15 ° after a st_Bady stat_ of rota_._ler_ted conditions.ll attempts to _ro_uc_ a _:pim f. It was possible with the modified airplan_ to pull _he stick all the way b_ck in tu_u_ing flight without do_veloping _Y tuucontrolle_-for motions either wJ.. T_e rudder trc_vel wc_s therefore limited to +_ 15 ° '_d _-_. The stall co111d not be reached because of the additional up-el. It was feu_d _. Returning to figure 1 it can be s_en that 7 final modifications were incorporated in the airplane in producing these characteri. Depre_slng the thrust axis._ics. In turning flight the instability associated with th_ complete stall in the original airplane _a:_ eesentLally ch.±on wa_ obtalned_ the airpl_ae would recover :'rom a spin after s_vorc. The a_ere.g._: up-elevator deflection required for a 3-point landing w_._.e s_me as from straight flight.f all motions acc<_mpazlying th_ stall was increased in turning flight somew_hat because of the effectively increased wing loaaing _mdF. the ai_l_me incorpor_ting all th_ changes except llmitiog the rudd_r tr_. ". with cLmr'%ctcri:_'tics ( ii ..s accompc-ny_T.m_ angl_ of b. but the z. The na:. The characteristics of t_.dy side. In turns the stick .tons.ng the stall w.zen_ral flyins in normal flight... The airp!&ue could not aoy setting_ of the ccntrols..cicnt to co_tez. _e resu_tin_ ] oncitudinal motion wss con_::idored !_ss dsn_crous them the lateral instability encountered ol.!ip .t_dder was still suff'.. _Jewer on or The .act the adverse ya._nk that could be obtained in a _te:.%l control dur__r_g the te_k_-off _s en=o_ntered.. the oriL:innl airpia_no.zof the all. prw%n: off..::'.:ec_usa of the !L_ted r_dder travel.Tr<_t!y rc6uced.'cs re!uteri to about half that of the crigll_ rdrpl_e .re L.is modified alrplauo mmy be (ill : v Motior._as held full back %_thout p_"oducinc a stall.of direction. were not materially al "erel be spun.. modifications to original NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS .- Three-view drawing showing airplane.3 RUDDER ORIGINAL TRAVEL MODIFIED ±15" ± 33 ° I I -3" -5" IN_I_:N _=_ELEVATOR WASHOUT _ I u _ _ ORIGINAL 36"- UP- TRAVEL MODIFIED 30.5" Fi_'ure I.E) .5" ± _ 28"-DOWN-10. 3 cd =: (3.) UP .30] POWER OFF ORIGINAL 20] ANGLE, I O-_ \\ ELEVATOR ON _, __ C.G. 27% DOWN I O-_ :50UP I. I I I "" "" I OFF 20- ELEVATOR ANGLE, I0DEG. 0 DOWN I0I I I I I 30 40 50 60 INDICATED AIRSPEED, 70 M.P.H 80 Figure 2.- Stick-fixed static longitudinal stability characteristics of the original and modified airplanes. NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS ELEVATOR ANGLE, DEG. UP 40TRIMMED FULL TAIL HEAVY 30- POWER OFF 20- ORIGINAL _ __ODIFIED.. OFF I0 ON DOWN 0 6 I I I I I I 20 24 C.G. 28 PERCENT, 32 M.A.C. 36 40 /'- Figure 3.- ',.L • Elevator angles above which instability existed for the original and modified airplanes. NATIONAL COMMITTEE ADVISORY FOg AERONAUTICS 49. FACTORS AFF_I_ING SP]I_SCING OF LIGHT I. Neihouse AIRPLA._IES Zy AzsPml be The _pin considored problem for ss twofold: of, or the personal-owner (a) elimination satisfactory or light al_plane _ny of the _nc_pient spin, f_:om, the _h_ll.v- and (b) el imi_mtion &_ _ _.l_pe _ s_in. Mr.• H_mt_r tt_rn_ng -L_ho, recovery has Just dSscussed elimination of _ho which appegrs dosire bl_ for preventing stall in inc_plent spins at//!ow altitudes. H._s discussion also included elimination o_' thejo_-pin of a light alto,_s by l_z.ting !he rudder t_avol. In thi_ connection, tests of numerous models in thu NACA spin t_r_nel have indicated tlmt the rudder is ve_ often the predomlrmnt con_,rol maintaining the s_in. haw_ • been conducted During the _m.r v;_ _ _ _a.._, spin inves+dgations in the NACA spin t_m_el on a_?proximete!y !_0 different mJl_terj _irplane designs to de,erm_.n., recove_ characteristics f_.om dew_loped spi_s. From the results of the investigations, it has been shown_ and repor b_d in reference l, tha_ an a_rplane In a f_l!_ d_,:velop,_d sp_n will r_cov_r rapidly _f suff_ci_,_nt _nd eff_ct_.ve centre! has been provided in its d_s[&_. Quite s few cf the designs tested h_d proportions cf mass and dimens!onal characteristics that s_,mulated those of airnlancs _n the personalowner catego_ a_nd aecord_ngl,v +,he results from t,_sts of military airplanes hav_. been applied to ligJat a_rplanes. In o_xl<_.r to provide effective control for satisfactory recovery from spins of 7-crser_%l-o_ner airp_lanes_ a criteri6n has b_en se_ ulo wh_ ch is an ext, ns_on of the on(_ previously r_ported .in refercne_: I and is sho_ in figure I. m_n._s-, criter._on combines the dame factors _reviously fotund i_Tportant in effect!rig recovery from the spin. As before, for _iffcr.?nt values of relative density, values of taildamping power factor have been plotted against a nond_menslonal ezDresslon f_r the difference in moments of in_rtia about the aicplane X- _Lnd Y-axes, and regions of satisfactory _tnd uns,_.tisfactory recoveries have been defined. Those regions _r0_,based _n results of sp_n !nw_stigatiens of many menials in the NACA spin turmel. Before discussiLg this fi6..c_r_ , more thoroughly, I sho_.Id first like to tell you a little about the, factors that have been plotted. Ta_!-dampi__ power factor (TDFF) is an Indica_lon of the effectiveness of the vertical tall In a spin. The m_thod _f computin_ it ha_ been fully explained in referc.ncc I and is illus+_,tcd attack in of figur e _. Dy 49 o, _h_ flow assuming a sp;,_nat at the tail beyond an average angle of _h _, hor.lzcntal surfaces / |. In order to detern_ne the values of relative density (_) typical of light airplanes. curves were obtained for airplanes having values of _ of 2 to _ and of _ to 9.a ds. minimum values of vertical-tail desig_ required instu_e satisfactory recovery from developed spins may be obtained for values of _ in the personal-owner airplane range. For wing spans of _9 to 40 feet and wing loadings from approxir_lately 6 to 2_ potu_ds per square foot. Airplanes having values of the plotted factors falling above their corresponding curve will recover satisfactorily whereas those having values falling below the curve will not r_cover satisfactori_yo For the relatively very li£_S]t airplanes v'_ry llttl_ vertical fin and rudder erea (_ = 2 to 5).Iy n_. a ra_oge of wing loa@_ings (W/S) has been plotted against-a range of spans (b) considered reprosentatlvo and the corresponding values of _ . that is.mping term and the second being a ruddercffectiveness term. plotted. The airplane relative density factor _ = _ is the ratio _aS. the first being.b 2 .roterms indicated. the corresponding values of _ for sea-level air c!enslty varied from approxii_ately 2 to 8. The tail-damping power factor is n_de up of the t_. wings in a along The uppermost curve shown in figur_ i was oreviously presented (reference l) for all airplanes having values o_" relative density up to 15.43 is assu2_d to diverge 15 ° at each end. divided by the airplane mass and span squared is an i_dication of how the n_ss is distributed in the airplane.oter positive direction indicate increased distribution of mass the wings. whether there is relatively more mass distributed along the or along the fuselage.. It nny be said to be an indication of the intere_ctlon of the inertia _d aerodjn_nic forces and mor_nts acting.b bet_reen the so-called density of the airplane (W/g_b) and the density of the surrounding air (p) in which the airplane is moving. the to fi_o difference i.have been indicated in figure 3 for sea-level density. P_eturning . the term plotted of inertia as the abscissa about the Xand IX in moments Y-axis. The rudder-effectiveness term depends on the unshielded rudder area and its distance from the center of Gravity. Increasing values of the par_. Thus from the dats. The wake thus defined indicates the rudder area which is shielded by the horizontal tail. The damping tectaldepends on the fixed area below the horizontal tail and its distance from the center of 6ravi_ of the airplane. is required to it appears insure that . For the present investigation. "eversal.... It can b. Hcrc too. but as a factor of safest..._ relative density of the eirp].gravity. factor _f ai_.'ecowr rapidly from the spin..c._._.w-_asod.-_rof gra_.r z_.irplanc re]atiw. however._.tj F center of _ravity L1 of airpl_e and centroid of area .? recover._nd 'Q acceleration TD._ .l(W/S) a. it _-. it _ccmes increasingly :'report. density pounds fee t factor .. especially if E_dder _ew_rsal is followed by moving the stick forwa_-_!. #IB.t X bod_v axis moment of inertia of airplane (fuselage axis) moment of :!nertia of alrplan(. As th. movementof the elevator do_ maybc of appreciable ass"stance in tcrm:_r_ting the spin. noted t_h_t as the distribution of massalong the wings is inc.tor. arcs. feet density of air. r Based on the requirements for v<_rt_cal tall design as indicated in fi_ire I..24 satisfactory recovery.ppearsthat proper considc_ntion of the factors involved will lead to an a._r_ne wh'ch will _.y..:.. it is felt that sufficiont vertical tail area shouia be provided to terminate the spin w_thout a_sist.. about about Y body axis mass of a! rplane of W .ng wing of airp]_ne.. slugs per cubic fo.e. square span.PF L tail-damping power disiance between of fixed area distance between RI c(_nt.lane and centroid feet p÷._.nes increases.qu_redto insure recovery by _&de.. at sea level) W S b 0 IX weight _._.ance of th.antto provJdo a sufficient TDPF to into:re satisfactor..arger values l of TDI'F are r_. . W.: . Lichtensteln. Anshal I. 1946.. i043. Tail-DesignRequirements for Satisfactory NACA TNNo. Jacob H. Neihouse.45 L2 R! R2 F distance between center area area of gravity and centroid of area R2 unshieldedrudder unshielded fixed area rudder below abovehorizcntal-tailwake below horizontal-tall tall wake horizontal REFERENCE 1. Philip Spin Recovery. andPepoon. _ _ 'HORIZONTAL .. for personal-owner type ... FACTOR = S(b/2)2x Method of computing tail-damping .Vertical-tail design requirements airplanes.. ELEVATOR /z (9 TDPF 0 t I I I -120 -80 --40 0 Ix-I Y rob2 40 80 120 xlO-4 Figure 1.'_.--I RUDDER RUDDER 8. TAIL WAKE I J I _.% L.. --L L J J COMMITTEE FORAERONAUTICS NATIONAL ADVISORY iRELATIVE TAIL-DAMPING POWER Figure 2.i (D xlO-6 0 .- /4 WIND 2 ASSUMED TO BE 45 ° FL 2 Ri L i + R 2 L2 S(b/2) power factor. c_ no °r-i i I 24- 20- 16- 12- 4 --- 8 2 1 I 0 29 I 30 I 31 I _ I 33 I _ I 35 I 36 I 37 I 38 I 39 I 40 SPAN Figure 3.Variation of relative-density factor with wing type airplanes. NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS loading and span for personal-owner d / / . 46 WING DESIGN \ . h corresponds to 2-percent c_ber on these airfoils. A comparison of an NACA 23012 with two of the newer low-drag airfoil sections is shown in figure 1.ds number of a personal. te_ts were made wlth each airfoil to determine the effect upon the aerodynamic characteris5ics of surface roughness. In order to provide the airplano designer with systematic aerodynamic data from which to choose low-drag airfoils suitable for different applications. The _<_dth of the low-drag range increases with increasing airfoil thickness ratio with only a very slight increase in the value of _the minimum drag coefficient. in that they were theoretically derived by potential-flow methods to have pressure distributions of a type permitting extensive laminar flow in the boundary layer and thus very low profile-drag coefficients. was doveloped here at the Langley L_boratory.se drag data are for airfoil models having perfectly smooth and fair surfaces and represent the optimum or idea]. Loftin. in addition. Low-drag airfoils differ from the older NACA h-and _-digit series airfoils. Two low-drag sections.and 60-percent chord.h. such as the NACA 2h12 and 23012.m here. Vatting the amount of cam_oer has no effect upon the value of the minimum drag but shifts the range of lift coefficients over which I_.ire 2 which shows drag results for a typical low-drag airfoil section (631-h12) of 12-percent thickness and Ooh design lift coefficient. drag.-owner-type airplane. a new type of airfoil section. The minimum drag coefficient of the low-drag section is approximately 9__percent lower than that of the older NACA 23012 sho_. The lift. . knoll as the NACA low-drag or 6-series section.vaising lift coefficient. and 9 million. thus it is possible that by a proper choice of camber the minim_n drag may be obtained at nearly any desired c. respectively. a series of approximately lO0 related lowdrag airfoils was derived and tested in a specialized t_o-dimensional wind tunnel permitting the attainment of full-scale Reynolds numb. the NACA 631-h12 and the NACA 661. DEVICES By Laurence K.dllion which corresponds roughly to the c_lising Re_mo].47 DEVELOPmeNT OF AIRFOILS AkD HIGH-LIFT Jr. are designed to permit laminar flow over the airfoil surfaces to 30. and.._ drag is obtained. An idea of the drag characteristics of these newer airfoil sections may be obtained from fig. 2/Z] _ / i During the course of the war yeai's. A design lift coefficient Of O. 6._'rs and having a turbulence level approaching that of free air. These data arc for a Reynolds n_mber of 6 r. and pitching--moment characteristics of all the airfoils were obtained at Reynolds numbers of 3.212 . drag characteristics m i¸-¸¸ . . and is extl_emely difficult to predict._ue of the minin_n drag _ith little difference between the drag of the low-drag airfoil and the older NACA 230]. The difference in the drag of the smooth m_d rough low-drag section amounts to about 36 horsepower.s corresponding to _%and or mud at the leading edge. the 2_-percent reduction in minimum profile drag coe:'fieient r_sulting fre_r_the us_ of a low-drag airfoil insbead of a conveniional one amounts to a saving of 12 horsepov'er at 200 miles p. The increments of maximum lift associabt_d vdth increasing c_ur$oer. the si_e of the wing and the Reynolds numberer.Reynolds number of a persorb_l-o%_:er-t>-pe airplane..n da_ indicative of the lift m_d dlag characteristics which mi_%t be exo. as tha_ of the older NACA ?. edges For an airplane ha_ring i%O square feet of "_ing area. _A_nebher or not these savings are significant will.cver. A detailed d_scussion of this subject is contained in _#_-c/+ ?N-i_l reference 1. These data are for a ReC-nolds nt_nber of 3 million v.[ih.th _ings which are too dirty or gritty to permit the attainlnent of extensive Lmminar flows and thus the indicated low drag coe_'f_7 cients. cu_-es shown fo_. The ma_imum lift coefficient is decreased and the effect of increasing_ airfoil thickness _'atio is less pronounced ?_heu the airfoil leading edges are rou._ that _th leading-edge roughn_s_ sufficient to c-'_usefully developed turbulent boundary layers. The exact size of the permissible roughness depends upon the type of roughness. both smooth and rough. An analysis of all _vailable dra_ data has shov. some muff ace Jrr_gularities may be tolerated in the range of R<fnolds m_r. The highest maxi'mtun ]ift coel'ficient corresponds to the 12-percent-thick section hsvin_ a O. In an effort to obta_. each of the lO0 airf_]_ was test_d with its leading edge sufficiently rough to cauae a £ully d_veloped turbulent layer over th_ airfoil surfaces.h design lift coefficient and as just about the same.ng up to 6 million or 8 million.48 that may be obtained. bers ex_end!. The variation of the maxim_n lift coefficient _'_th airfoil thickThe upper indicate ness ratio and camber is sho_n in figur_ 3 for a ntu_oer of NACA lowdrag airfoil sections._r_ hour. the two airfoils with rough leading " _ar_e increases in the va. of course. however.hich corresponds to the landing .2. Surface ur_fairne_s or waviness re_ult_ug from manufacturing inaccuracies may also prevent the attainment of the expected low drag coefficients._ctod for the worst conditions of _surface roughne:.3012 sho_wn here for comparison. the minim_m_ drag coefficient is relatively inseusitive to ai_'foi] shape aI_ increases with airfoil thJck_ess. to be relatively insensitive to airfoil - . Many airplanes operate _f. depend upon the percentage t_m5 the wing drag is of the total alrp]an¢ dra_. B. appear suri'ace condition.o_. '_iott_d flaps show theft th(i decrement in maximum lift coefficient caused by leading-edge roughness is the uamc for airfoils both with an4 w_ thout flaps.r than those obtained with split flaps. the airfoil wit.l... and slotted flaps and for a 12-percent-thick low-drag airfoil with a double slotted flap. the airfoil best .._b]e in reference With the use of this report. i_dlcating that the slotte_ flap is much better from the standpoint of short take-off run. however. In any case. show that. definite advantages are associated vr[th their' use.mit_ d for a given th_ for 2.re h for low-drag airfoils of 16-percent thickness equiDped w_'._o-d_mc.9.rfoils with split flaps as well as ..ected the drag coefficients obtai_-:d for airfoils equioped with slotted flaw arc _uch iow_..5-p<:rcent-chord leading-edge slat and a single bounda_j-la:fer suction slat locat_d at l_O-percent chord. .tior_ of the tolerances to which the wing is to be manufactured and the conditions u_.0 can probably be obtaint_d with these high-lift dew'ices on airfoils of 18-percent.... the characteristics of low-drag airfoils ar_: no worsc than those of conventional airfoils and. A co_Jplete presentation and analysis of acrodynnmic data for lO0 related low-drag airfoils as well as ne_':t. includ_ careful co_sider_..th the low-drag airfoils.i]:e.. therefore. the maxlmtuu lift coefficients of the airfoils __th split flaps increased up to thic]-.h doub?_e slotted flap had a maximum lift of 3.. The results obtained for ad.'_sio-u_-I data the I..th_ cknes_._e that maximum lift coefficients Some data ar'e available -_. althougb the highest .der _hich the airplane is expec5ed to operate.and %-digit zcri_cs airfoils at.nesses of 18 or 20 percent.._ obtained for thickness ratios of approximately 12 percent. The problem of developing good high lift devices for various airfoils was the subject of considerabLre%arch before tbc war with the older t_'qoes of airfoils and during the past few years this work has been continued _..'.maximum lifts for plain airfoils wer.. . if sufficient care is taken w:ith the surface condition.. Section maximum lift coefficients are shown in figu.80 have been measured in this thickness r_mge. It is 2erhaps of interest that w__'th flaps d(_f].th plain. Naximum lift coefficients as higl_ as 2.nd_c_. split._.Thich "_ " "_ well over _. These data. The predicted characteristics of airfoil sections _../?J L_ 9 The lift and drag data which h_ve b_en presen+ed for _irfoils both smooth and rough indicate that the aerodynaf_c characteristics of airplane wings are _tron_ly influenced by surface condition.now _w. With a ]. Each of the airfoils in the systematic series _as tested _rlth a 2e-percent-chord splib flap deflected 60 °.ngs must. some data for .._ ap}_lied to airplane .. Summa_of A_rfoil Data.r Re)uuolds n_. : S_n_mary .. Some of the data includedinthe s_1_na:_Y ai rfo_. coeffici:2nt coeffici_ent pounds pound_ (d/qc) (_/qc) pressure. Wing Sections.. John H.: . L5C05. /' / SY_ABOLS cd c_ d q c section section section dynamic airfoil section drag lift drag.-! report are now being e_Lended to lew. 19h5. lift.. I_!.mlo'u'sfor use by the designer of small airplanes havi:_[Z low landing speeds. of Drag Characteristics of 191:6. Subscript: max ma xi_/m TB Quinn. chord. Jr. Louis NACA ACR _4o. Coz_truction Jr.. von Doe_hoff.7 be selected. Albert E._O airplane may be selected.. _ _- Practical- 2o Abbott. So. and Stivers. NACA TN No . Some of the data included in the summary ms. Ira H. feet per square foot po_unds feet t airfoil thickness. 412_/_ SMOOTH COMMITTEE FOg AERONAUTICS ol -1.4 .008_ .4 0 CI .016/ . .. 1.024- . .032- .2 I 1 I I I I -. i \ NATIONAL ADVISOI_Y ..8 -.Comparison with the 661 -212 airfoil sections of shape of two NACA low-drag NACA 23012 airfoil section.--NAGA Figure 1.°3 _O f NACA 23012 NAGA 631-412 <.01?__ .OO4631.8 low-drag edges.6 air- Figure 2.2 1.020ROUG .028- .Drag characteristics foils with both smooth and of NACA rough leading and conventional 1% = 6 x 10 6. 04 ._.4- 0 Figure 4.8- / _/_--'_ / .0-" O 1.4CIMA x 2.20 t/c Figure 3.4 .thickness having Variation of maximum ratio and camber for both smooth and rough section NACA leading lift coefficient with airfoil 63-series low-drag @trolls edges.2 Ctma x .6- 25012 _/_-A ..08 .-_. _ 1. 4.Maximum section drag airfoils having various lift coefficients obtained with types of high-lift devices.4 C1 I I I I I 0 .01.\ } sMooT.2.63.61.0%..2 ...22.." . 3..'-_.-._ _ t/ .. R = 3 x 10_J.8n f_. NACA low- . NATIONAL ADVISORY COMNITTEE FOg AERONAUTICS .82.2. ] ROUGH DESIGN .16 .12 . stalling of a wing can be determined . lift distributions and the A re9ort. Wh_-n the m_. The calculations indicate sepzratlon ov. A comparison _s given in figure 1 of calculated. now in the process of presents a method for calculating fo_'ce and moment chezcacteristics the of j wings using nonline.._ver.. a sl_. I is g. as would be expected._ and the section .ubst_. These.!_ft c::rves separate is belle_<:d to be an indication of the possible rate of stall progression.: It should-be remembered that the methods presented are subject to the limStations of lifting llne theory. the section lift curves . The point of Initial. portion of th_ wing. _. The difference between the curves is the marg.ght disturbsnce m_y produce a stall ow_r a lar$_._n between the actual lift coefficient and the stalllng. while those calculated usin_ linear sec_Aon-lift data are in agreement only over the linear portion of the curve. Although the distributions of figure o show a sm__ll margin between . Stalling should beg-'. The _. STALLING S._ximum lift.are not linear and hence more refined m_thods of calculation •al.from ctu_ves show_ng ths distribution of sect._ section-lift _l:_t_.5.section llft curv.MAXEMUM By LIFT Harold AND H.reberg The problem of prgdictlng the m2. together with the distribution of the maximum lift coefficients of the individual sections.progr_gsslon for th_ same wing describ.iatedby the results of tuft obse_-vations shown _n the upp_-.nt._ndencies. The calculations are given using bet/% nonlinear and l_nesr section-lift data.r part of fi_uro 2._on lift coefficient along the span..central and inboard portions of the wing sem_span with no evld_nce of ark¥ t_p-stall_ng t._& _n fLgur.ximum lift coefficient and the point of initial stall img._rgin _s sm'_ll..curve. facts arc _. At h.lgh angles of attack.rthe . necessary-..ate at which the section lift and section maximum . An example of calculations to determine th_ point of "initial stalling and the possible rate of • stall .. e publication (reference l)._ft coefficient of the sections. ho%'.:.n at the points of tang_ncy of th<_. of a wins from available twodimensional airfoil section charactgristlcs can be most readily solved by the use of lift_i_g-lJne theory. _a%d experimental lift results for a w_n_ of aspect ratio 8 and t_per rat_o 2.lven in figure _ together with some. The close agreement shown in figt_ ! should not be expected for wings of low aspect ratio or large sweep. Solutions to this problem have been prossnt @ by numerous writers using linear section llft dats. experimental results. The lift curves celculated using nonlinear section-lift data are in close agreement with the experimental results over th_ entire range of lift coefficients.l. by wing-surf. and by propeller operati on. were used. Full-scale w_ndtunnel tests of two airplan_s showedthat Lncreases of between lO and 15 percent in the maximumlift coefficients of those alrolanes were obtained when all protuberances were faired and when all points of air leakage on the wings were sealed. say. 230-series sections. possibly indicating a more rapid spread of separation coupled with a sharper drop in lift past CIT_a than are shown x in f!gures 1 and 2. Analyses of wind-. P more rapid sprea< of separation. flow separation from the inner parts of the wing will generally cause a nose-downPitching momentaud thereby flncrease the longitudinal stability near the stall. Wing surface roughness and air Icaksge through a wlng should be avoided wherever possible inasmuchas their effects on_xlm_n lift ar_ detrimental.ness. it was shownthat methodsare c availeble for predicting with sufficient engineering accuracy the maximum lift coefficient and the point of initial stallin_ of the wing alone.tunnel and flight data relating to the chang_)sin the stalling characterist_'cs of airpls_ues resulting from nartial-span fle_p def_!. thus providing a stall warning. Furthermore. The loss in maximum lift coefflc_ent caused by small 8_eas of separation _J_ound fuselage the studnacelles will g_nerally be compensatedby the added lift on the fuselage and nacelles unless exccptionally poor wing-fuselage and wing-nacelle Junctures are incorporated on the _Lirplane. The ouestion nowarises as to how the results of these calculations are modified by the addition of a fuselage or nacelles to a wing.oction have shownthat the results may b__divided into four main effects: . edin_ discussion.'_ce roughn_ss and air leakage through a wing_ by partial-span flap deflection.52 the section lift and section maximum lift curves except at the wing tip. This effect may tend to improve the handling characteristics of an air_)Lane n_ar the stall inasmuchas the wake from the stalled roglons may cause tail buffeting or a loss of control effectiv_. _t sho_uldbe noted that the calculations were madefor a wing employing 4h-series sections which possess gradual stalling characteristics. and a sharper drop in lift p_st CLmax would probably have been measuredfif sections having abrupt stalls. The addition of a fuselage or nacelles to a wing will generally introduce small local areas of sep_iration near the wing center section. In the pre. . an_.th w elliptical chord distribution and sprea. of .p. The atall originates about _.'er_ses i.: t.... For all tbcoe ai._ximum lift coefficient was measured with flaps dcflectsd than with flaps retracted. stall originates at the wing tips and p-_oE..._pered w!n_ (to. -.ling ch .io (t_'.._ and_ progc.i)er ratio_ ]._ts_.. third air:)... WJ... flaps retracted._ deflection on ths .._ -___ wing trailing at i edge near the root .ly to the angle for i..aae. a wiug with low taoei'_ . in each case.rA.r':_ two. bi_s of an sirp]..'..'% _s characteristic of its the particular exception o!an for.!._)..ace has _._rdl_:):"tionsof the wings.'-.ccea..ca_ The smalla.. O:-..e hi. It is worth ._t.. The lift curves correspo.rd from the wing tip for the _.._e._.._ having un_..Jn i_.....-.._. For tb_ wing with Io_¢ taper ratio the stall o_.r an. ""'_ _'_ _. c-. ... ior_v_ is EJv..53 (i) An increase in upwash and effective angle the outer unflapped parts of the wings the tendency for tip stalling... the stall progression for each win.._.rjp+:ar-_...!th h:.nts shewing _he ef_tects of partial-span l'l:_]. tendency wings.._ses outboar.of tai.. No small..._.tt_ck.. (3) A tendency to cause a more than in usually measured (4) An increase envelops in stall the tail sudden loss in lift at the stall for the unflapped wing._l._a_el mea_J_L_om_.__ . !'i_ure 3.. anotLar haa. _o_oa wh_n_ the fourtil effect tends to improve the l_o. _ .._]. :._la_e Junctures.ing:.. s Full-.ration behind the /..¢ deflection tended to dela_ separation over t..a..scalc-t.e ai. to delay separation over the of attack over thereby increasin_ (2) A inner parts of the \z..urx'ed at a hi_her angle of attack with flaps deflected th-_.e of ate ._ inb'. ntion!ng exceDtionally poor that flight tests showed stalling characteristics...bf. and warning provided the flap wake at moderato to high angles of attack... a more abrupt loss in lift past the _... . .._ a ._-. iuiti_._-_c+io_..saio:as of three a. For the wing w.ane near the stall.2_:1.nwi':.... the i_f_o.th increasing angle of s.--o '.__ea of attack for earth stall Drogression _hown in fig_ro 3 correspond roughly to the a_gle for initial small_ng while the hi_her angles of attack correspond rougl..._oard _.ding to the stall progressions of the three airplanes noted __gu_e 3 showed that. .ng.! outsoard with increasing angle of a_tao_..i:_.. ' chord distribution........nes.:.'a'. ._h taper ratio.O percent of the semispan inbo._:'_u_..iginates at .t _d. w:.] .th flaps retracted.. _l:e _...._ _." nacelles .. The first three of these effects tend to produce -?cor stalling chara ._pl ..._. with wiEe\y dlffer_. _ ._' -___.. _.."and the.rpla-.inc on'.. and for the other .1_..-c. .. with of small localized areas of sops.._t pla_.._r_dat -_]ie wing-fa:.._:.......rc[ parts of th_ win._.l sbml..o: :. airplv.. This effect was very pronomlced in the case of the airplane with win6_ of elliptical chord c!istribution..o have .'.:?: .._...ugs prog:'.:e_: fla.s both inboard an._.bh 17.has a wing " with "__".._.._ximum lift.. ti_is airplane t. sharp leading edges will generally cause . 4)_ whereas.. which would be sufficient to cause roiling instabil_ty in landing. central sharp leading edges. This featttre.-on l_. incr_asing camber from root to tip. ':_ith the propeller operating at a t. however. Lateral instability and the consequent wing-dropping tendency is perhaps the most serious feature of the stall. The use of airfoil sections at the wing tips of moderate thickness (above about 12 percent) and of fairly hi6<h tambour (about 3 or 4 9orcent) have been shown to improve the behavior of airplanes having bad tip. the rolli_g-mom.th0ds . An asymmetric_l stall pattern is thus produced.2.hrust coefficient of 0. and l.:_ading-edge tip slats are discussed in d. the airpla_le rolling-moment coefficient w._. Both theory and experiment have shown that high taper is conducive to tip stalling.sessentially independent of angle of attack (see fig.ding with about 1/4-rated power applied. several methods have been devised for moving the location of the initial stall inboa_rd. The use of thin s_ctions (9 percent and below) at the wing tips should be avoided wherever possible sspeclally on wings having taper ratios gr_:ator than _oout 2:]. Above this _ugle. which include _ashout. which con-esponds to a power. a sharp increase in the rollingmoment co_Jfflciont occllrred. the stall progression with angle of att_ck ls typical of that for a low-taper-ratlo wing and is fairly symm. Measurements of the variation of rolling-moment coefficient with angle of attack showed that. the wing sections behind the upgoing propeller blades stalled at a considerably lower angle of attack than the wing sections behind the downgolng propeller blades.2. Flight measurements showed that this airplane devaloped a serlo_s left-wing dropping tendency during power-on landings._r only one wing panel to lessen the c_nsequ_nces of sewrc asymmetrical stalling. with th(_ propeller operating at a thrust coefficient of 0. that the a loss in maximum use of lift. It should be noted. Owing to the loss in aileron effectiveness and damping in roll usually assocJat_._tr:cal on both sides of the plane of symmetry.snt coeff:Icient increased gradually with angle of attack up to the angle for maximum lift._tail in r. Th<_se m<. can genGrally b<_ avoided by considerations in the preliminary d_sign stages of an alrplana._ference 2. Sharp leading edges haw_ also been used to induce s_paratJen ow.stalling tendencies.54 The effects of propeller operation on the stalling characteristic_ of an airplane c_m be illustrated by the full-scaletunnel measurements shown in figlrce 4. with the propeller remowd. With the propeller removed. which is mainly caused by wing tip and asymmetric stalling.d with wing-tip stall. _. j . it is felt .Spercent of slats th. flap deflection.lllng ch'_ractcris tics. extending over 75. Throe cor ditions were inw_stigat_ d: (I) (2) Airplane Airplane without with slats extending over 53. Flight tests were made on on_ airpl_ue to investl_ate the off oct of the spanwiso _xtcnt of a lcadlng-cdge s!_t. Th_se calctulatcd wing char_._on.e semispan gave tl_e airplane the ability to ret'_in a h_terally level attitud_ throughout the stall. It is concluded from these results that.ny _'ight_ng tendency.d.8 percent of W_th no slats.8 percent of t/.lstallin(_ of a given wing._r._.. howe_. by proD_.ctcristic_ _y b<_. for this a_rplane.y rapidly over the outbo_rd portions of the w_n 6 and th_ _irplan_ _. I wo_id like to reiterate that methods are available for calculating the _'ff_octs of the wing @_omctry _nd cf the s0anwise lift distribution on th_ max_oum lift and the point of initi'-._n m_.._neral uo._as_d d_'ag if the sl'_t is f_xod _nd added mechanical complications if the slat _s r_tractable appear to l_mit the use of the leadir_. Increasing the extent of the slat to 75. stall_n_ progressed w a.53 Experimental data have shown the leading-edge slat _o be an effective _ans for reducing the tendency for _icontrolled motions of an airplane _n a stall. however. These and other d_ta indicate that thJ !argo slat span required.. by win_. The results are presented in figure 5 as time histories of th_ stall ._ sem_spsn (3) AlL-plane with slats the semispan.. The extent of those effects can be most readily evaluat_ from analyses of the _vailable data on maximum lift e_d stalli_z research.lat_ons which were not m.-fuselage and w_ng-n_celle interference._ith a icading.malned in a banked attitude.'_rch has recently be.a_.d_ available for g.l_r opc_ati_.. F__nally. In this connection a su_m_ary of available full-scalo-t_mncl data r___o_t_ng to maxim_n lift _d stalling r_s_._nc r_. _k_ i _ In cenc]us.-ed_e slat as a means for obtaining satJsfactory st_.. together with the disadvantag¢_s of !ncr. the airplane exhibited obJoctionablc lon_itudina! osc. radically changed. by partial-sp.5 percent of tJ_ s_m_span the airplane initially rolled to th_ r_ght but th_ roll vas halted by the slat so that th_ a_rp]. rolled rapidly to the rig_ht with no evldonc_ of _. th_ opt_imt_m extent of leading-_dgc slat would bc about 65 percent of the semisp_._ (refere_ze 3). and by peer wing-surfaces conditions._rog_ession and rolling velocity during the st'_ll._d_o slat e_tending over 53._asured for the other two cond itiol_s..i]. . propeller di.mb-er l_ft ar_z!e of R c_ _eyno!ds sectio_ coefficlent C_ rol!inc-moment coefficient (qb_. density pounds 7 slugs feet .56 that-_sufficient expe_.._er cubic per second. feet foot-po_.and with chs.re now available to show theft by carefully selecting the proper combinations of w_ng (_eometry with _¢ashcut . square '_ttack. feet Subscript: max m_tx_mum . wing span._ds Tc e__ct_vo .imental data s.) L b roilin 6 mcment._p satisfactory airplu_ne stalling charac-. D2 ) T D V D thrust._ea._neter.. _gMBOLS CL _ \ qo q S dynamic winc pressure. bounds feet de_!_es per sq_mre foot a. nges _n ai_fo[l camb_r end airfoil thickness from root to t... foot a'_rspeed. thrust coefficient _0_ . n_. teristics may be obtained without the use of auxiliary devices such as central sharp leading edges and leading-edge slats. _ni D!ngeld..g/&'j__ Ch_ts R_.rist:'cs of Airp]ane... H_.940.Scal_._ Theory Using Non_-Lir_eaz" S_ction Lift Data.:. 3.eparat_on) 9.)t:!n@_ of T':_pered Wings._ar_. LSC2h. and N_ely. }Robert H._7 _S I. Sweb. R_char_ C. (Paper [n o_.mgley Full . 1_45._lllng Characf_. No.dn. : De." Method for Calcuiatin_ Wing _. to the A. 703_ ]. NACA ACt1 No.].. H._rLst_cs by I ift_nT-L_n. Sive!!s. Je_mes C.s.ct. H.ro!d. SouL_. i' . : S_mary of M_as_u_ements in L_. _h:mle! of Maximum Lift Coeff__cients and Pt..rg. NACA I_:_. and Stalling Ander*on 3 R F. 6/" .0 charac- UNSTEADY _ STALLED 1..6 SEMISPAN . NACA 4416...4.8- . for wing ..6- EXPERIMENT--\.32 I I x t06 --1 4 8 a.04 TAPER RATIO.. -i_" /_I_CALCU LATER_ OQ I.O- • CL .2 18. wing lift NACA 4412.6Cl max 1.4. a. b..5 R -.- Calculated and experimental stalling characteristics described in figure I. deg 11. 2..4 cI 1.2- /_ .4 .O 1...1.Q CALCULATED (LINEAR) 1.. 12 DEG. Comparison of experimental Root section. I I ASPECT RATIO.4.O NATIONAL COMMITTEE ADVISOI_Y FOI_ AERONAUTICS FiEure 2.0- . 16 20 Figure I. 8.o") 1.8 I.2- \ / I' I | I i l I | I I I ! I c t (a -18.teristics. and calculated tip section.4% 15.2 FRACTION . . ---J 13.._ 13. 20 NATIONAL COMMITTEE ADVISORY FOg AERONAUTICS Figure 4.5 17.1 16.9 61 12.8 FLAPS 16.1 PROPELLER OFF T c'0..2 11.. 15.9 19. 1.9 16.. Airplanes 16. .5 RETRACTED 15.2 18.6..:.3 12.- Full-scale tunnel measurements showing the effect of propeller operation on stalling characteristics.3 EXTENDED on stall plan progression form.4.h0 (D _J O3 (:Z.5 I. DEG 12.._.9 10.Effect of flap deflection of different tunnel.9 • oo_ I I " I I Tc-°2 " I 0 I0 e. ! FLAPS Figure 3. DEG.1 of airplanes tested in having untwisted wings the Langley full-scale .2 16. SEC. on 20 Measurements to investigate of leading-edge effect flaps stall- ing behavior extended and of spanwise extent throttle closed.5 i \ 5 II 18 I 6 16 ROLLING VELOCITY RAD/SEC. NATIONAL COMMITTEE FOIl ADVISORY AERONAUTICS .5% SPAN SLATS 75.O f-. slat. 0 I 0 TIME. 20 TIME. I \ 17.b. Figure 5.8% SPAN SLATS EG.> CQ NO SLATS 55. 20 0 TIME. " 5 SEC. the I0 SEC. in flight I0 SEC.- I I I I I I I0 SEG.4 C> (]. _S f .( PROPR_LT.. the ideal cf]i'icicncy and th_ id_ai . Th_ upper curve in the fi6-aro (th0 solid line) ils the sc_nu :_s giv. hlghcr spreads.powcr for operation with thu propell_r moun:_od on a lOO-}:ors_pow_r engine. (Se. Th_ m_thod provides a breakdown of th_ propeller power losses _nd thus allows t. Althou_ considerations of i_.T _ A John L. Experimental for _hi=h verifications of this test da_ are available method for specific have been obtained applicaover the tions entire oI_erating r_ngm._ reforenc_gs Several of propellor i to 6._. the propell_r di_ocer is. This method can be used _ des i.g a propeller for a zi_on airplane _nstal!a- tion based. EFFICIENCY Crigler A m_thod of selocti_.ercent and a 20-foot propell._or propeller is to absorb lO0 horsepower at 30 miles p_r ho_._am.a5 can b_ obta_nod in the take-off rang_.'ate the breakdown of operating conditions. Figt_r_ _ shows the _v_ilc_bl_ t_Lrust hors_._rs. of co.ic_ncy at tx_½_-off speeds are therefore unavoidable with prmctlc_l propell. The tnr_e curw_s represent three propeller d±_ue_srs. Lot us consider next the conditio_z we will consider a 6-foot-d_amot<_r at flying oro o_lL_r spe._rs _bsorblng leO horsepower. this m_im_._ an optimum propeller for any operating condition sa_d will accur-_tc!y predict the propeller perform_ic_. Th!s figure covers the take-off speed range only.h.) prepared a ranG_ to illust._r to about 80 percent. while a 12-foot-dismoter propell_r would !ncroas_ the m_xLmum posslbl_ efficiency to 65 [. It is seen that the propo_[l_r diam_._ds._n possibl._ter fixes the maxim_u possible eff!clency _b.ty in the slipst'_._ best compromise of th_se losse_ to be made throu6hout 5he en_r_ operatln_ r_g.ECTiO_._ c£ficlency is o_nl_ 48 percent._] _fficioncy favor large di_et_rs in the t_k_-cf_ ' rangu. charts i_£_vebeen power losses for Figttro 1 sho_s the variation of the ideal propoll_r efficiency with speed for propell.-J '5: PROP_7_I_q By S_ZT. In this as _. 6 fe_t.!ations a_%d on o_cgorim®ntal verificasions _ro _lvail&blo. Serious losm_s of _ff. Sinc_ _ lO0-horsepowcr _ngino was used. P_iblications on the methods of _moretic51 calc_._n in figt_ro 1 but has b_en extend_d i<. upo=_ theoretical calculations and exp_. This ideal efficiency is caiculat_d for minimum axial en_r_D _ losses alon_ _nd n_glocts the drag of th_ blad_ sections as w_ll as the ro[atiom_l v_loc_.ypical of case presen_ practice.a'. a_l 20 fe_t.rimonta$ da_:a has been developed.Arso_ limited by prachJcal consld_rations such as prope3_l_r wei_'_t 'and clearance. if a 6-foot-di_n_z. 12 fe_t. . Changes in _°_k_of' _hosc vari:... on the pr<:T_!lor hub and tip sections in ord._l]..= _c]-:_ncy over th_ s_oed r_ngo.."operation w_ th a 6-f..rvo _:_d tim c_._r to give minimu/.._. i porocnt lc. The .'....or _ffici<_ncy is o..6o in the fi_...oen this ct.ud . soctlon thickness.._- power for operution wibh th_ same prc)p_].gatod as o:lo L' b _ -u ch_ng<_s in the propo!icr lo[_..oti.in solidiLy _s t3p]ca! of bh_ . sho'_.... The!._ _.o._Jd _ii'inu tn_ rpm _ecreases.oo__d by _h3 o._gron_int:_ined the in. . Sinc_ the fixed pitch ]?re. 7' ._ller was oz_ly sli_. th._ decr.. shows the calc_.:nlO. Th..i.. . Calculations woro ma&u or th._._muz load distribution at 120 miles per hour :_._%¢o tr._/'foctsof ttiri'oil socbion. In the icw-sp_._ dosi_ condition._ho ongip_c d_croasod. resulting in a doore_se in t]..nc_:..S ].iLated _._r with ideal thrust po_.. the loss dlw to th_ proi'.ius i_. the 0...._ith fixed pitch..'ffect e of 01. from r <o_..)fleet (-f a_!.o wotd.edline _..ncy for the cx_r(. miL_s o_r hor_-" to 130 m_/. The :i.il_ _h'a..ac:_:'_.tcL_ <._ nora_powor of '.. th. If t....:asJ in the brake horsepower. i_c._. p_.blcs .l onur_ loc_.:.17c<ntov_.£ch distribution.:i..n_<_cc.r wuald absorb more po_:er at constant rpm as the a!l'sgccd _o d.._ ._'.dian_etG__.:r 0ban the bb" saving of Th_ the axial dashod line energy. two.r_crcasu..th. oaad solidity on b}]o propcl!.. This cur_o is for vai':_ab2.'v<..h%r:.. w... pe]. form and • . ' " ..T_l?-_r with Clark-Y sections ard ro_d b!ad-o sh_k:3.'mt braks horsepower of lOO over 0he enb_ro spc_:d raui_e..:.:..hr_._..a blade.£_s.c<_. for the 6-foot propoll.o_sos.r _ne -_ange of' sp_o_s...ivos l..'n axi_.fix.. showud (krLy r [_ppr<:o:kabl<_cham:]_ in the <hrust horsopow<_r over [ ._r _._ hub an& tip s.r_]uJ i_orc._ion. pl:.[C"::'thf_<.d ._d pitch prop _l!or to doter- '\4.[ encr_' in c_us_d by tho decrease in load..r is about...l_d cm-v_ but _.nd.ron_ of thos..rhioh '..idictribu.r the pr'>[_oller w.]_o or<k_r of 1 to 2 l_orc_nt...powur dcv<:-. b]'_ I_rofil_ drac of the bls_u_ and axial and rotaSion_.'.t horso pow.0345 per blade.._ o...7 r.-blado !_r_._ly !].rlbuti._rur_ h_. ' " r. bo!h aru p _ tc_ d_s_.. .h. Extr..d< .h._" !0 ]_ml'c:. i_c!uding the round blade sha. The difference _n power losses b_tw._s for a const.. solidity pron...Ally e_uscd a change in th_ bhr_.2s booaus_ the propel!or l<.>r of aoproximat_.9 changes.n rosulL in -.the rotaU!o_:'_l cn_r$// loss..p_ratkon iJaving a_ opti.L_r b!ado h_s an activity f_otor of 90 and the Jolidi_y at.J".t prcfilo drag of th.2_%: cna_C_'.Ld by the _rop.u_]'._ pr..)s p.._ ¢_mr<_ ]. to but bho thrust hors<. Small differences i_.._roL_. The major el!florence betw.'_=_s calcul_tud for a similar blade bu{_ with a 50 p_rcont increase . _&._ . Tb_ t.ho load on th._L!._ang%_s_n p.tnfornl._d..es_ two curw_:_ Ls _.-:3.o incrd_sc..)!l_or [..ust bors.:_inc../ rain._r hour unless rad]......adin_[ is not th_ ideal as giw_n tim mol_bntum theory...or is (liv)d<_d into three parts..:'r_asc in rotat_ona! cn_rg%v a_d the tip losses would b.g of the blade sections.p_ilor _ oiTioi_... the tot:_l d!ff_ronoe in _ropoll.i __>_.r for oh.l_r at: t!:._o_.tly increased buoat.._n th.._ro.'nv.r<ter of _.ob ._ £rouL.. and an additional axial onerg_v l(._s_ are pr_seut but.. l)r__k._ctions -./Lu_t. of th_ two lqw_r c_r¢os _.._m-ust horsepower fo_.h<_tn_._ 0._st hors_pow.mdashed.po_r output.n '.l_.b. ment In thrust horsw._ller that is d.a_ volcsit_.ira8 r_i'.tn the 1. the f.:. variabl_-.. ef V/nD a_l small chs.oower.wb of a Oltult body such as an NACA cowling and 4.:bo_t opero_ting rang_:. tu-iml In th_ tak_-of_ rang_ the drag loss .culabions.v_loci'By dist._ op_rat_d in a low _l.61 reduction these gives off in the t]_must horso-.llor s_ction. ght air Diaries can be ro_llzcd by further r_search to obtain _mf}rov_ propoll¢. os r_sult .. _ If this nonullzfor_nlt_ .m_r _n the take- range. t_ird of the ra:']lus of" t... al'.r &!mat-_ly 70.3r optimum wh.dra6 iuss._ of apsrox. If good._d to opor_t.at_d separately from th_ &rag losses and a prop.5') haw a lift-&rag .I[.zt o.n the prop.. of:I:'ic!_._ll_r may be consi&er_}d as two supaz_.£& froF_ler for £r.nB8 of _bout 0.9o-stroam op®ratio_ is no long._rccnt of 8 the total power but h_ro a6e.'r off'io. Fr. It is thus seen that of all variables.. In the dew}lopm_nt of +.. 0 c_ .[_%toly 1 perccnC in the propcll<._ region.&.ribuBion _h. airfoil saction'_ extended in cl...:rm_n. Th¢ of the drag loss for a_. ctit. section lift-drag ' .rel}eratin_ condition can be dot.gain in propeller efficiency for .ado sections plotted against V/nD___ for x a l.'. In the calculations 5b_ change in velocity frrm_ the fre_-:3tr_am velocity simply rosa!is in a c]lange in tho e:l.._ tak. ang.nj_s in tm_ s_ction lift-draz retic du_} to cilan_c_s in tho a'._.ro_mferonc_: such as is tl]<}cas_ wh. Figure 3 shows th_ ..oso to . For _x_Mnpl. a given operating condition is so mounted i__ front of a blunt body that.& thaB no stlboua:_tl_l ._.ra_io of _tp[.rat_s in a field oi' nomtniform a2.-:._n pluGci ill a field of nontuliform axial valoclty. It _y be conciv.._ of values of the ratio of ]._ in._lemon_._%1. s_Ippos_ a prop.e V/z_D of the prop.._ncrcases to ab...?-P.}ly.m.._rag us_% in propolLsr ca]. fe.r . One pruduc!ng lift alone and the otbor producing dr%g alone. We shall no_¢ consider brieflj th_ cffoct of bo<y Interference on the p_opell':_r officlency.. efficiency loss due to t_o p_:)file drag of the b].io of th_ sect.Lgmrs the imporUuace " .siLnl an an optimum dasJgn. since the propeller usually op_.ions i rosuit s in a chan_/._te propellers.. It is Lmportant that ohis . propellvr " .._2±_._ improve.he propeller theory _J_a inducted losses are tr_.sri into accottnt in the propol!er &os!gr." aection8.nitch operation is bh_ orii_v ch. Typical sections ope'P_tt_n_ at lift co.. chanzos A c_ngo of +iO perc._. r officionc2.'enoy..at any appr_clabl.an6_ t..3 s_cn that _.n. _'"._llor is mounted !n fro... "or{ant. which is the .'_ll<:rdc.± b-.3sJ..lens operating at a lift-drag ratio of 70 in th:.=il cbancus in the Is ""_ ill __. d total power loss &Lie tc dra_ is about 3 percent for sect. is cc!_stant aro_n_& bim_ cJ.in i5 i.rfoil or SoG'BioII thickness are r:._ i3 ]_uo_m bh_} _.:£f]ct b_ included in bhu prop..ffect may b..the hub those i_ler s.._L'-_tlv(.ift to profile drag. This shift of the loatd distribution is not CONFIDENTLIL ._ctions would be owrloadod with a restultant decrease in propo!lP.'. the In_.. _nco_.'_d ff climb.cy"by about 23 p_rc-ont.!_t.80.)c._pollc'_'wero d.:.d by fla_?.or.'.[ sin[_l_-sl._signed to operate in the region of high v..t_p..i t.or that can only be d_tel'mine4[ by hosts on tho particular' arrang_n:_nt._n_-. Thls loss in officiency could b_ c_us_._conditions [.ry good for th_ 0 ° flap scttins.nsist._nt 9ropoiher (:ffj. A g.._7 _" " o± _c_.o 0_0 For tl_e sa_p.2 _radion_ with .oo clc:_ _ in fron_.on [:_ such that somesections arc c_l%pletely unloaded.y ca:._ v_.. of a wing bu_: %h_ vu!..m and tho velocity d!_tribution in the nropeller nlane is _hnown the res_I/. Defloc.%round the circl/mx_.. Althou_ the propeller w_.h_ efficiency by abowt 8 p._rc_nt at V/_/D = 0.ions.afficic._be very accurate!y com.streaa_ operation. Mo.srence: _Id the propeller can be u_zdo optimum just as well as for frse-..ions t.'.:_ pr..tJlt-c..62 serious.c_.Jr oi.-_neral arra/_ge_nt of tho tos t mo_i. Figure }_ shows rest<Its of t.[.77 and power c<.__ :sums Wfp_-_of flow with t.h_ split fl_q_ r_duc_d the _.j-_ :3__... cause_ nomuna'ormi-cy of a.the pro o_lh.o the 5!on of a split flap.n. If tho airfoil characteristics 5re kno.r _ffici_ncy and tests are a]ways required to de_ormln-_ the efficiency for a particular location. clo_e up b_d_ind a wing the effect c_m_noz be inclu&_d in the calculations..o n_zdc over CL tango of b!ad_ an_Tl.. Th.opre&ict wilat happens _o .ocit.h@ resulzing loss can occur ii' the prope)..: l_tndin_.puts_ for operati.orm axial volc.J_ would stall if o_.u of a h:_6h-sp..i'ulsive ufi'icli_ncy of a propoll. .. ncy i_._d of tht_ installa- and Tests wer. 85 percent at a V/_D of 0. If the nonuniformihy in axial velocity cuts across the propeller disk as in the case of a propeller mo_nt._d..ed. of no_utnif.f£. g<.f t.s moturtod clost. _Ln_.oionc._]. original model c._t[.y J..xiu) _ir flow.'.[. C . for low V/nD operation In%les:_ is enough to it stall someof the sections or ur_\es:_tho voL:oJty dictributi.:d..._lificat.zrry t!_ load alternately thr'.-u_h tho regions of high and low volecity witL_out stalling is a _._.-.. however.L_s of attack to be .._ mcdul usud was a cor_s+.:l_nt.h_r the propoller can c.'._s_ flap d_fL_c!. It :s difficult _.i.}_u. tant_ .n "n a £'iel_.ol to duI:.orated at tho same rpm and pitch oetLing in a low vu!oc't2 region._ing the sincl_-slott_d flap _o 40 ° rod a_. t.!._d bombur. up b_hlnd -_ l'_rg_ nacelle it should 0_: noted that the proptuls!vo . T]._r is mounts.licatlr_Z the arr'-o-ngcm_nt _t 5bo cen'_or nac_l].l_±'ioc t_d f]..lis sh':._Ling ohind b a wln_: with a ._tt_dflap du!_.ars or arc/ hod.l.lord wing with _acello an...oity_ i_rovide_.os_s that we!-o co_luctod in the NACA propellor-r:_s_tirch turu_. this nonuzllformity is cenc.. For _ns_al.autore_lin t... Wh.ormino _h_J pr. _dge._n such through will be tho low conditions velocity ._zlst a_d region the blade s t_%ll'_on passing efi'icioncy tl'.63 distance is much sharper in front of the wing t.-a _.h_nd _.)pei!er is mo_mted very close 'to the ieadtlng . cf the '.e_ent f" a&_ancc-d'amcter or _:. _.han bc.al.ect ratio of propeller of blade elemont lift-drag propeller r ratio tip r_dlus ele_nt <_f blad_ clo1_nt (r/_) radius ._oti.. _d_B 0 IS D D L propeller drag l_ft dia_ ter blade element for !nfi1_:itu as_.lement effioi_ncy efficiency leas due to drag .%.._ing.nles-a the p_-. there is ices likelihood of getting into troub]. to any locaticn :c radial l'l propellor axial rotational speed V velocity of propeller ratio propcller n propeller _].e]os_es in propell_r vez'y serious. llorbert W. from George W.. F._:_ for Propeller Efficioncy. O!auort. (Berlin). W. 440-_'_5. pp. of Springer Aerodynamic Theory. NACA ACR No. John L. . Julius Vol. pp. vol. H. Data.: No. Soc. NACA Theory of Screw Propoll_rs.64 RE_I_ENCES I. A. 6. 3. IV ally. NACA ACR. Determining 5. Six-_ and _gh. _!B04. J ob_ L.: Airplane Propellers.. ed. (London). no. 1941. L4129. Propeller 712. : Comparison of Calculated and Experimental ?. : Char t.. 2. Sydney: On th® Vortex Prec.. 79 ° . Stickle. 1935. John Rep. July 1942. 1944.. and ?alkin.n Aercdy1_mic Considerations." i" Propeller Characteristics for Fo_-. ser. Durand.. Ar_lysls John L. 1944. L. Crig!er. and Talkin.:_ll_r Sel_cti. Go!dstein. 19c9. Roy. from 4. Crig!er. : Prop. NACA ACR No.mrimental and Crigler. Herbert W.-Blad:_ S±ngle-Rota_ing PropelLers.. 169-360. 123. O April 6. Ex. L. Crigler. ....- Variation with propeller..... 80 90 IO0 VELOCITY......... T .. I10 I 120 1 130 Figure 2. / • / _ _ __. FT. 1-i 60 50- 40 0 I0 20 30 40 M PH efficiency NATIONAL 50 60 VELOCITY._- HORSEPOWER I / / /// . MPH of thrust horsepower _]-foot-diameter I .I 00- b_ 90- PROPELLER DIAMETER.oo i I I . !......._'_ _-'_=------------_-'... .} . ! N... 0 801DEAL EFFICIENCY 70 H .- 70 _1/.. velociLy fi_r :t ..... . Figure I....................Variation of ideal propeller with velociLy. ADVISORY 100-] COMMITTEE FOg AERONkUTICS PROPELLER DIAMETER... 40 50 I 60 w 70 135 I I... .8 0" FLAP_.02- I I I I I I I I 0 I 2 :3 4 5 6 7 8 Figure 3o- Element efficiency loss due to drag for a range of values of L/D. 4O 6O .6 V_nD .© EO • o . ""\ ///'_----_40° .-.8 1..06' I1..4 -.0 ADVISOI_Y COMMITTEE FOIlAERONAUTICS efficiency at NATIONAL Figure 4.0 ._ II 15. .t.080 I --1 .6// L40° SLOTTED SPLIT FLA.048O 120 .- Effect of flap deflection on propulsive constant take-off power.4 0 I " I I Op=.oo'---4 SPLIT FLAP ..2 .12- L/D 25 .-I .. E2 .. FLAP LOTTED _ FLAP J . <4q • 3' 16.I0. _d m_t that the Dredcminat:'ng so_'rce of noise in present . dlsc_issi:_n will thcre.: the rlct:":_m'_uln6 factor in locating %h9 near s. l_ yea. Some of this %_erk is repel:ted in refe_ence !.'..e..'o_to±' the p_[sscn_:nrs of the ai_21.Ve. the chief _.].'. CA it is po_nt. I.:'n _nd_'st"u ' _5s in hi.__. n_is_ w_th special roferenze to light alrpl6nec is given in a r_._l._i_zs-._ m. was discon+. to_y airo!a.:_o.._ interest of t_'e :_vi:v[. ho_r_._ore TI_OEY OF PROPELIP_ NOISE _ae propel!or noise theor. that in or. work on pr_.rpl_es.3.%_Lul_.irport _.e:" nels. T_V. _'lq[uw_._l_!)(.<_c'.]it.Ii.q_ .:'pl._)-.:_ca_ion: . sc !t ma. .'actic'_]. Th_ to propelle_ present nois_.e.uos _und _t was f_:lt th_.xp_ns'!.I.._o h.e:r noi_..'s ago. Work on thi_ _n'ob!em wa_: started abou _.or..s pvoblcr_ bec_u.-.loeb r_. oe a_r[_!_./ _h ___[:no !se o f" !i ?']_t:-i _.l was developed in }_ussJ_t by Gutln. _.r_]:'. wi r[i?_i_.._ speed of sound in conf:_rml'k._ • . TTis h_Tpothcsis in s_mLo.t .[s._y +.b!.ecs something is done about the propeller._er no'f.d in th_..on pro6r_/m.e nor'.L]J.._-'r _nd possibly the _i.ceference 2.. _. At "[:hatt!mo.'_i:.inued ablaut _ ?. h?_d been foun_i and oon_i-_med.!y. vici-_iL-.ia'!a_-_"cr'_.-a._'. A].vcr.re silenced it is believed that engines can else be s_le_ce:l by conventional be restricted _6%hcz.ir.'._.. _fhen prepcJ.T_0_ By Arthur N0_SE A.6.ler to e!i_fJn&te p[t':.ea...ears ago because a _Tenerc_ theory of prcpel].ez l_iooz'atory of the EAC_. /-'_ .hat the sto'_dy aerody_cmic i'crc_ (n the blares (thrust _hn(ltorque) are imp:_rtod be the :.irports prcvLded for [n the new Vederal _.n].y with the:: clan_ic .._..[m'_%-'. . for milltzry a_.fts the _ propeller and that _'ilcncing the engine is u_cle_s t._. A gener%! review of t}_e prob!em. • of' sunil airports._RODUCTION (]ii_ r_:e prcble'_u of airplane noise i s not a new one here at th_ Ls_g3.[' [lO_CC_ Or materially rodv_ze it_ a radical redesign of the p r De]. Regier lq.ck has h e_sn c one <_ rned pri_zs.)_:'_']l] e._vlng in the.ion of airplsne noise will gceat!y increase the oe__J._.]. Thn second theo.lors "'.<':s_ai7.s : s hear_ by p<.v b. pr. _Rccen+.cen*._r 8rod reefs sho_.%_:bc_n roz_:m<d &_ the reqve_: of D_".m.it aud propagated through space at th. The CAA is intero_f-_. the sl.. C.?.oole ].d. p.r"A. -h:.herefore receives a square wave impulse whose ma_itud._d int_i?-'atlng the co_. Some work on w:_rtex noise is rope:steal in reference 3.rcssure for a_ point at .s .. but arc oscillatory dist_. There is enothc]" ecru'co of propeller noise duo t..9_ificant clean up to be achieved on prop._mi]_a:"to that of wind blowing through a forest or cf waw.ir in the propeller disk a-t a He ass_s that no forces it.nl output of' the propeller Is _ucrcascd above tbat given by the t'uc. There is no si.(!am_-nt..mdlF sources the b!-_dc as are those considered j.%de up of th.ne the air act on the air resolves a repres_nta%ive radius.. Ic: conc_.swh._.]._rom all "Jy f the particles or_.n the blade are caused to vary by nonm_ifo_7._llers tu_til tl-e tip speed has been reduccd sufficiently so that t. Gut!n noise on the other hanl has only frc:q_cne. In orler to simplify the mathe_mut._n . These sou-_ce__ are not st.'.._in..s breaking on a beach.yhave ach[_'ved this _inly b:[ reducing the vort_:x noir..e thr_st of the blade.s on tip load conditions and flow condit_. T_:. For lo_... For o_8mplc.ages oz" _.. For nonnY_ propellers operating a+._in:_<s_ the so_..'_"y..-ry. The basic . flow fic!dm _uducd by fus_.hc _utin noise is b_lcw the vortex noise level.c.b.ad . The air particle i. The s_ . The sound 7.u& output.omd..ch ar_. _ho .tex noise m_y beco.1 blade passa_ frequency._._..._ distance from the propeller is ¢.aud as observed by the i_ stoner is s.._. household f.n the blade. . Th-_ frequency from these.u_.r receives a sharp blow _.O.%ns in recent year.o_.._n% r_anufacturer_ have developed f.r of 0..!_:_r ._il! bc dosiiB_at_.tr_n ra_g[nz from a few h_udred cycles per second to several thousand.leo_.5 thee <_. It is becm_se the a#..rn(:d only with the steady a_.e_ _.i'.res_::i.onics of the fm%danental.9._ sc_.. the vo:.tdi_t_.66 laws for sovnd Gutin co_si_ors propo_at..the vortices en shed from the propeller..atl. If the propeL.h!ch de0en4.ons ..Q in Ibis dJ scussion._.. If the forc.e is _qual to tb...'b_auces.e to st_/ody air forces ".<rc!_ .uld-_.d in. .me large compared to the Gutln noise.of the fu_a.-t!p-sgee& i'_ns_ however.lon. vort_x noises are small compared to the Outin noise...mbe. Th. a small sc_ction of _.x disturbances is almost a continuous spe. until the blade reaches pressure equal to the pressure on the blade for the ti_e that the blade passes through the section of air rudder consideration..n the G_tin th(. t:Ip speeds above a Mach m..cill_tory vcrt. This _qaare wave is resolved into its Fourier coefficients or harmo_ic_.r forces on the blade.:.multlpl:_s ..v'_ section of the !_..lcs.r. "_ t. Thr.b.described by t_e rcl..3 to reduce _'eatly the sov. At that tf." Gat_n soo_.j _ .hen the blade pas:ms it that propeller noise is m._ fundamental frequency (revolutions per second times nt_nbor of blades) and all integral ha__. ditional loss is a function of th. It in well la_own that doubling the spool of a propeller increases the pc-wcr consumption by a factor of 23 o:r 8..foling tllo tip speed hhe hors_p<.er{_ radiat_d has been incroo. tip _pcod it increases so_nud output Thus me& be seen that by do_. If there is no di.of The ratio of the Gutln noise to w.s noise down to the vortex level.:. propeller hs_s increased by a factor of 8.-.nity _rithout regard to the absolute ma. The Gut_n so_md shown _n the third co]. to be "_all for short distsnres.wcr to the. This a_.s much sound ener_?_v to bc radiated per hersep_'.ed.r._ propeller.tip speed is l-educed sufficiently to bring the Gut. but the sound ez._nn w%r:_es _s the tenth the power of the tip by speed.Tillbe silenced to about the level of traffic noise. ho:_evsr.lo_ing expression. the pressu_:e will be 20 DB _e__ or _0 DB at a dlst_r._ed. It is because the solmd energ¢_ w_rios approximately as the tenth po_er of the tip spc.co of 300 feet.l. The vortex tip speed.propeller.or to t!%¢.6._iJ "_TrT_wo .._os are given for twice the tie s_._Itudo of the quantity. that tip speed is the most tu_portant factor in reducing propeller noise._sipation or less of sound ener_'y in the atmocphere. the seinedpressure varies inversely as the d__tanc_. Table I gives the effect of tip" speed on the sound of th. DIoo_o. It is Imo-_m. Thus if the sound preserve is 100 DB at a distance of _0 feet from the so_n-ce. In the second llnc the relatlvo values of the qua_tit.5 pow¢_. The change iu sound pressure is given in decibels by the fo]. the propellel _ _._]: (lO_h/8 = 128) it is seen that doubling the t_p sp_ed c_ases lP8 times s._ sound frequency -and a_ospheric hu_idity an_l has not been _ccursto]y determined. _dg.henco_do_'bl_ng a "iactor of _10 or the loeb. If the relative Gutin sound is divided by the relative power to the p:_'opc]._rtcx nolso is the .. d2 where _ is the ratio of the d__stances. ss For large distance accotuut must be tauten of the additional reduction of so_.u%dpr:_sc_ure due to the _Issipation in the atmosphere. In the f_rst l_ne of t/_e te_ble all quantities a_o given _ val_e of _.._lON OF .sed by a factor of 109_4. noise given in cehmm h varies as the 5._v_%. This fla_re !_ shown in the secured colurn. 3om_._nd experiment on tho 2-b!ado pro_. The good acree:m._.."._ n_. w._r..ad(: b by tl_.nt dir-_crepancy of about 2 decibels which may be duo t. !o_-::_rlthm In subseqi_ent f. .le±_s. ca!i'bration. Hence the vortex noise rmn be __-.:.cr indtc.9. but may become the :_./o.:. It may be seen that doubling the t:_. :-_e.r+..ig1_yes the &ec.%bfo of Gub!n . power.gh. iro re_.y i_ not considered u._...68 given in the last column.:!ues for the 2-blade rm! the 7-b!.un v._.pe]lorr_ of many blades and (2) to (_ : _ ._.ad= :_o .%_'e given in figare I..go of tip Mash n_.d be useless."' e_:.t_._ / d_sc... l./nbers from 0.s givc _.. so'_A reading will be referred level of 10 -16 for watts per of heazdng SeCO_. There i-: a const_.9. It is seen.i level as c..__.e prone]._.rcdominate :_c.i.': n propeller wou/._[eterm.t'..bor but becomes _ _ _-_i a'b tile low . o[' L]Ic .deastu_ements m _-s. At the h..rtrx ?.'.e r. .i propeller so_mded to v..lgh tip speeds the sotm':l is almost a pure note equal to the blade p':.!co_lated by the Cutin theory.. but were used because they we're available f.lerc were :o.ne vori.. the: h1_. F'0.i...b_is..suremcnts.. the The v(...nt between theory .bel to the con_acnly a¢c_..e NACA ! an5 7-b!a._ to 0...ise f-'om such -.oller cunfi_n-'a- tion.._p ..:. no he._e_z'ch lh.. 0 soee&_ l_c._.. .. In the second lind third co].'}l ":_I s(:tmd sI.ubJn propeller sound..4..::n be obt:_Incd for th(. It is i0 tim. It sho%_ll bc emphasized that for a given p:rcp.._. th_q/st. The decibel is a ten_ cor.pted base eentimets_ which is the threshold at a freq:.7_s the leg_.:ency of I000 cycl<..]. or ZO decibels.!..'e..rce of noise at _e.: the nc. how(::v. no]re 4:o voLPtc. _ th::.also g_od for the 7-blade propeller at the high tip Maeh n. worse v.¢ tip sloee&s.Der.e:" tip _pe_ds.._ricu'z listcn.m_#_.'_-_.ve. Ti_._._l ear pcwer rai.ecen5ly some . and torque. the Gutin theory predicts the sound pressure. T1_..... ller tests The . _.l.._ pj-ope].0% designed for tltis problem.Irum obtained from the 7-blade propeller confirms this belief.o_ some l_igh-speed prope.) microphone.... on-t.lstic C. =.o.r_...7.-.[ is .eated that the C sound ener_ ratio of I02_ is equ__i to ZO DB cr decJ.blade propel!o:" over the..t:'.uv_-. The dashed l_nc..cm_.ex noise level for p:.I:.. frequencies and its ha_monics.*_.s per square on 2- P. Tho a_grcccm_nt b_tween oxpei'iment and theory is gee& for the 2-.results o:C these tests . to be due to in the Outin theory..rf_ues of the sou. tip speed.olid lin..._'_n it is ind.:or..h_ theoreticeJ[ v. Thu_ I-'-_. particu!r_r test cond....n.]._y 1.x noise by a . This p.rihhm ef the of 1024 is '].._..ll.undth. propeller tested is _s good as c..ou_-'_t]".._Y_ tile hi._nomly used in sotu_d me. _. _ for .o_ give the s experiment_.it_ons of power and tip spee'.oif:e whi._. _o_of the tests were (I) to .<..beru. The agreement botw_:en exDerlment ..-.-. o_-> eo. that on the.:_r._ :.'._:y lo'... At the low tip speeds th_ vortex noise drowns out or masks the char_cter. 7 tlp Mach nt_uber. energy in the higher harmonics is loss then for the lo_er harmDnics the total sound prossu. etc.ry. the so_uld pressu_:_ is down by lO decibels. 6. The lowest line gives the th. the Outdo pulses are less for the 7-blade propeller thc_u for the 2. This subject will be discussed la_er. a reduction of _l _ecibels.tip spe_d_ the measured values exceeded the mlnlm_za v. _ i/ I I . t_mcs the propeller rotation speed._'esholdof hearing. The 2. Thi_ means that.a. 28.opo_ate&. h.s indicated when the c_. S__nco the so_n0._e _s less for the 7-blade propeller The gain for the 7-blade propeller at high tip speeds is not as great a. This is nct intended as a practicnl app]_ication but serve_ to demonstrate that the sound pressure c_m be c_reatly reluced by operating at lower tip speed and by using a greater nt_her Of blades._'E_en from roferenc.blonde propeller absorbs 59 horsepower _nd the 7-blade propeller absorbs 97 horsepower. Th_ 7-blades propeller radiates frequencies of 7._ h and is considered as a standerd by accoustlcal workers. times propeller rotation speed.mnan ear. etc.:'_u_o only ha1_on!cs of the bl_Tde _assage are p_'. It is desi_aatod zero decibels The loudness is defined as the loudness of a pure tone of I000 cycles. 21. Thuo_ the 2-blade propeller radiates sotmd at frequencie_ ef 2.buted over more blades with the result that the lea. The characteristics cf the human ear are given in fign_re 3. It may be noted that the CuD propeller operating at a cruising speed of 2100 rpm is absorbing _0 he 'sepower :und radiating sound energy at a level of 107 decibels as measured at a distance of 30 feet f_Dm the _ropeller. Fi©_re 2 shows two propellers _h_ch _ere tested on .69 7-blade propeller o_erst_ng at Io_. This chart is t. the horizontal scale the so_md frequency.blad_ propeller.l •per blade is reduced. •Tha wrevious discussion has dealt !_r_-marilywith the sotmd pressures-_ithout regard to the frequencies or the loudness of the sound as heard by the h. _ero is another cain for the 7-blade propeller bs_. 14. I_rther resgarch _y be usef1_! _n _d_cing the vortex noise on thls propel ler. 8. _ne solid lines give the loudness v_!u_s as obtained from listening tests ma_e _:ith a laa'ge n_nber of obse'_ers. The vertical Scale gives the sound intensity or prossur-e _n decibels._acteristics of the }nm_u ear _'e consi_lered. The 7-blade propeller having a _naller diameter _s absorbing 48 horsepower at the same rpm and the sou_zl energy _adiated is 86 decibels.%n electric motor. The reason for thi_: _s that th_:_ thrust has been distr_.._lues predicted by the Otltln the_. The horselJo_¢erto the propellers is indicated for both propellers operating at 0. In spite _f the fact that the 7-blade propeller is absorbing nea1_lY twice the po_:er. If the ai. It has b_en sho_n that the s_md p_:essv.O c2cles. The characteristics of the humeri esr oleo explain why helicopters axe inaudible at a dists._equency the same as a helicopter. _. to have %he same loud_ess requires that the so_a_l pressu_e be reduced to 60 decibels. If the frequency spect_um is raised.ro can be reduced by dec!_easing tip spe_d . ess is obtained.nd the fisst three harmonics fall in a frequency b. It is necessary to make the combination of' sound pressure Intonc_ity and frequency such that the airplane will be on the same loudness level. In order to mak@ "_he airplane as quiet gs the hel#. Th. The airplane requires or_ly abDut one-tenth ._lane propeller has a frequency of lO0 cycles per second. simplification of the problem because pr<>peller noise is not a pure tone. which is a freq_emcy range approaching the lo_z frequency cut-off for the ea. ho_._ever.a _ater solu_d lutens_ty maw frequencies for a given loudness level._. the int_nsity must be increased from zero decibels to 60 decibels in ordel _ to keep the tone audible. the loudness may not be reduced _uch because this inc_:eases the frequency. be sound intensity below to the lower frequencles_ tel. If the _ound pressure is reduced by adding blades at a given tip spewed and diameter. There is no basic rea_qon vhy _.1._nd by increas__ng the m_ber of blades.uce of a few h_mdred feet._'.co_ter. Thus if a hel. therefore the so_md pressure for t_e airpl_me is much le_ for a given tip speed. for any lO0 decibels the ea_: is much less sensitive therefo. _f the frequency is re. The Ytu%damental scund freq_ency for a helicopter _. becaus_ of the characteristics of the human ear. Thus.orated at the low This fact must be consid_. If the frequency of the tone is reduced._ve the ssale loudn_s_.rod intensity of 80 decibels -_t a given d. it is net necessary to make the so_u_d f.istsnce at a frequency of 30 cycles per second it _:!ll have _ loudness of only 40 decibels._f the thrust. a creater intensity is required _o g. ._e.7(> therefore. airplanes cannot be made to o_e_ate as quietly as helicopters.iuced from lO00 cycles to 'i.red in a design of a propello.__ is en o_e_. In general._. to illustrate the essential i_>atures of the problem. contour as the hel:[copte_'. decreasing the tip speed aloe decreases the frequency of the sound as _. The difficulty lies in bringing the frequency of the _xirplane propeller do_._. the loudness rosy be increased even though the sound presstu'e has been reduced. For a given propeller. thereby placing the noise in a frequency range in _ich the ear is more sensi tire.copter has a so._xd bet_een 15 _d 60 cycles per second._e ex_nple serves.zell as the pressure _'_d th_cfore an additional decrease in lo_r. the loudness levels and the intensity levels coincide at lO00 cycles._ to the rauge in which t_e ea_' is insensitive. _ise from Rotating 4. Z. no. Noise Reduction with Reference Ar%h_" A.:_her. : Vortex N_. 2.ion. 747. i14:5. No_se D_e to To_2m..71 REFER _NCES I. Acco1_._. Jo_n'. 3. vol. and Theodorsen. F!et.3. PP. Arth'. an& Demin$ A. Deming.. A. Oct.No. Theodore%: ond Relier. : The Or*blem of to Li_ht Airplanes. Definition. Am. . 195. : Loudnes_J. NACA C TI'. : Thrus% NACA TN Propeller No. 1946. Stcwe!l.m' F. 82--i0_. Harvey. l_otation 1940. %'. Its Mes_s_rement and Calc_la¢. F. end _[unson. Soc. E. 5. 0 Figure I. compared . NATIONAL COMMITTEE ADVISORY FOR A/PONAUTICS 130 59 HP TO /0 P.. propeller as a b..S) 5.5 ° PITCH'--.O (I) pd Gut i n Sound (. ./' 9O //--/ . PROPELLER TO 97 HP Z//.'..ted Vortex Noise (T. 7ATBLADE@ PITCH PROP..-i I . 12 THEORY BLADE GUTINS 7 8O / d I l l I I I / / 0 . t ] n u [{ound Sound Power Tip Speed {T../." .Relative sound energy for a two-blade function of tip speed.'_crgy I i'.- Test data for two.. .Table rE) .6/a /s IlO 16.S) I0 Sound Energy R_d ia.and seven-blade propeller with Outln theory..' 2 BLADE // 0.5 P(_er to Pr ope 1 ]er VoFLt?X So_md 9l.9 I00 GUTIN'S /_'" // A"--- / THEORY--------/-._ / .6 NUMBER .4 MACH .. 45 128 23 TOTAL SOUND IN DB PRESS.S) to Propeller (T.S) 3 Energy Rad fated (Gut in ) (T..8 1.2 TIP .OP L ER// 120 2 AT BLADE PROP. I I i i I 1 1024 2 8 (30db ) diff._. . Sound Tip Math NO..DB COMMITTEE FOR AERONAUTICS 120J __ ____jl20 I001 _ 6O 4O t/'i' I I I 1 I 2O I00 FREQUENCY Figure 3. 701.57 12. .5 _' 4...' Pressure Figure 2.. ill.'_ ...45 I R.39 21..': [I It.. P.... H P .M...- Test conditions for sound recordings. tl 1' _ourl(l ..M.... Pitch ._... -2100 . Pitch ..... ADVISORY NATIONAL INTENSITY LEVEL ..8 ° 40 i07 db NACA IITGH-:.A[)E...UF_.. .. }_rcs_.PB]_:D PI.SOUND b_ PRESSURE AT 30 FEET CUB PROPELLER Sensenieh NO. ...._1o0 . Tip t_ch No...P... ..IN 500 I000 PER 5000 I0000 CYCLES SECOND Loudness-level contours. _l!er noise (refe-cncc !)... require some form of pitch. and the necessity fer h'_ving a starter on the cn_.....) On this f'igurc is also ._ geared bec::_se. try to for_ct the _:os_'... It is ger_erally recognined that the worst noise so.i_ear. control._........ a level of /E: TM 7h inches in I00 decibels.pber of 0._ in take-off perfor_aance. ho_ever.svzer to the problem...ecds. weight. It .ffectiw'.. propeller _!ach number is taken at the 0... a 2-_lade propeller... the nL_nber of to be the only Everyone i=.ERS _........ tho:t of reducing pro>oiler "diameter... before.. dia_aeter.. At the present tim_... It has been shown that propeller noise may be reduced by reducing propeller section speeds and by increasing blades. c&nnot be tolerated because of the penalty inc_r. but it seems that the propeller is much moz._ comfiguration.....e bu__ness c i.i. hut necessary... in t.... .. at 25_0 revolutions per minute..._±... It is powered b _r a !25-horsepo'_'er _n$ine rated ....... reducing the dimmeter is undesirable from take-off perfo_:_ance..:_ involved and ju_t see _¢hat hapi_ns _hcn we attempt to d. crul:_cs at 120..._ direct--driw.. For this purpose.:_ to add up to increased cost._ IN THE DESIGN OF PROPELI._cihels c_d at an effective propeller Mach nu.. Let us... It _...PROBLE_._m complications.6 (the c... as m....... Th_s propeller produ. increased mainter_sncb. and it is assumed that 5he increased _rithout radically changing our present L_t us now see what this 8-blade design 8-blsd...ed at I00 d..±Ing...:. _. Voge!ey FOR NCISE REDUCTION Sy Arthur that The increa_ed publicity noise reduction is now being given to sirplan_... and complexity.. this ap!_roach appears a_.::_ign a less noisy propeller for a typical persona'.: worrying about what Lhis mc_ns to the light airplane. The present 2-bisde o_rating point is loca_. The _...... .ne :!_:Lacc it will...... It is relatively easy to quiet the engine by muf_....._n_...ill mean ge_r reduction since the only other alternative for reducing propeller se=tion si.. be d!Iflculs to crank the engine through _ . The airpianc chosen has a top speed of 130 _ni!es per hour... 7 ._rms of second mesns reduction....:_s sound at The design chosen to reduce the noise is an propeller.8 radiu_.e indicates '_: '_ not merely desirable....._rc_:3 are the engine and the propeller... . The diameter has been kept constant..e difficult to handle.. airp]....hol.. fig:ire 1 has bec_n prepare<_ from fi_Dzre IO of Theodorsen and Regier's report on prop...... for the moment....[oreblades me_. and takes off at 5_.ntioned a consideration of diameter cannot be airpla_... ..et<r.. {]omparisons have be_._rodyrmm._:..a]. acco'_ding to t_Jcory..n made both cs a fix:.ance v:it!i tht. "_- f ... it s]_ould be remembered._ctdrive propeller...te!y _ pcrocr. :n-o-)_:ller_ the take-off dLstancc is reduced approximatelF 22 pmrc... By going from <'-... discussion noise One Lhat _._'<. 2-b]....t i:: real J '-' Thi'.d._.::.t [onal speeds...... prop_'_ ..n unchang. f_'om an a. Ci_. um speed and cruis:h_6 speed remain prsctical!y unchanged but the take-off dis!. problem. Con.n_.:mai....i_tionai contro!li_!e prope!!e_s.. the sound level should be reduced to something ]... in his paper....... Crig]er. u :: ........:d-pitch and as a controllable propeller.... lqgure p_:rticuiar case The ne:d_ two figures _ui!! show the relative performance of the 8-b!ade geared prepc!!cr as compares to tb.._ ......_.d cr'. has sho_..0.' 'r call b_ - ..lpitch .. " ..ie coritrollab]. It is seen that the :_uxiF.....l to d.es_ than _O de..tudy has shot'.e can be accepted.nr<_ v:j...'l 8-blade i propeller is the only possible shows that many other solutions has just been taken as a first answer to the are possible.'.t low rot:.. Duol...._.e_ will colu.:...cted from L...' b< bui].:m: the new desiz:_ condition._ origins. number has becn reduced approximately _0 percent and.' to a 2-blade direct-drive cc. it is w.73 sho.._in'.a.._l_..re again..or_w...ut if a contro!]ahit.s.r'c t_cw the 8-blade propeller_ if made contro!l_ab].. From _: a_rod_m_ic standpoint_ the requirement of quieter operation _ntroduoe's no additional problem.._d prop<<iler prepe!!er_ l_oth opersting in Figure 2 compares 5he performance v.e propeller.. ...._....iDh the original 2-b!ade direct-drive fixed pitch.ua sound It should level not of approxinmtely be ir/'erred from 60 this decib_!s.. but if it can be t._d.. The effective l'"acl'.ucti<_n in t:._ f]. }!_...y..o.ro.. of the _-bi_. h_ _' re.t:_ vo_'tc... the take-off performance can or.... it is hop¢d th:..lybe increased by varying prope!le_ _ _itch. These comparisons are ba_cd on calculations using available theory and data._d... oler:_._...ng pitch control is remo_:'ed. an <xtreme cnse the solmd level will sti]! rcm_}in below ih..Gie..._tby ha]<iug fuel..m Dlmt with a giv. : ._g].2-b!ade dir.:D.:n propeller diameter..ntrollable.em when the propeller o_er_:tes a..3 shows this comparison.rising speeds r._ may nDt be willin C to accept this reduction ]n taRe-off perfoi".:eoff distance of approxim.ic :_. brief :.. however _ that the present thuory n. _igure . ma<im. a £.. st_p..._m '_. Ccnsequent]..t.t'..de gmar.....na:ice.._.orcent.xc.. _ quieter fixed-pitci_ propeller ms?..:)ry will be optinistic....:'ect-driv....de di..n that..trt..._ndp.D _ some ll_cp.sc :in takeoff d istanc.. Nevertheless....ng_ng to eu 6-bl.....t.edthen the nec-_ssJty for providJ._ desirable C maxim_....-_equently_ the sound reductior_ to be cxp._i:.. 8-blade propeller was been b_croased about 8 p...: noise which becomes a proL].::ib. l... _[r. 7_ built then a sm_ll _mprovement be obtained. in t_ke-off performance can actually As a mattel- of interest, figure I: shows approximate_Ly wh:.>,t the 6-blade propeller would look like. The blades e_'c more or less conventional, the hub has been nmde l_rge enough to contain any pitchchange mechanism yet not se large as to increase the cn_Lne cocl:Lng problem. I would like n_.,¢ to go back and review some of the problems mentioned earlier and present some randoI._ thoughts Woout these problezJs. Gear reduction means increased ,t,vight.but, siuce gear reduction is already necetsary to reduce noise, then one can, l_rbaps, reduce the weight penalty by redesigning the engine to operate at higher revolutions per minute -:_th a resultant increase in horsepower per pou_nd. At first glance an 8-blade propel!or wo!z!d appear to l'e four times as hea_uT as a 2-o].ade pro_.i]_r, but, b_.cause the rotational speed has been reduced, the c_ntrifugs! forces hay[; been reduced, and, because th<_ number of blades has been increased, the thrust and torque loading p_r b?.ade has been reduced. Consequently, the blade stresses hnve be._._.n r_duced enormously and thi_'_ shouT_d make po_sible the production of considerably lighter blades. A brief stress aualysis, considering on!7 thrust, torque, a_id centri]'ugal force, indicates that almost any nmteria! th:_:t will hold its shape would b_ structurally satisfactory. This ma_ open an en_,ire!y new field of blade fabrication. The design of a pitch-change m_.chanism for an 8-blsde prope]ler would, at f._.rst,also appear to be vers" comp!ic'Lted, hov_ever, J.fit is rcme._bered that a 2-position i,ropel!er is a goo.[ s_sti.tute for a co_Zinuously w.riable--pitch propeller (at least under the conditions of light airplane operation) then the problem is considerebly simplified. Furthermore, because of the low centrifagal forces and io-_ rotational :!_pet_ds the design for controllability is stall luvLher simplified° !t appears possible th_,t sn extremely simple 2-position m_cha_J_m coul_ be built. Producing a less noisy prop_ller is a very tough proposition, but it is hoped that this study has shogun that if quieter propellers are demanded, one can, by caref_l consJdcrabJon of thei. peculJarities_ _ reduce the penalties thst must be paid. ................. .... " ...... [: .......................................... 75 PEFEZEWUE I. Tneodcrsen-. Theodore, and Eeier, Noise Reduction _.._itheference r No, 1145, 1946. A_thtu- A. : The Problem of to Lig_It Airplanes. NACA TN CII: ,y. , t_ PRESENT SOUND C) 03 h9 © OPERATING CONDITION PRESSURE, DECIBELS I00 t80 BLADES2 6O 40DESIGN 20CONDITION / I I I I I I 0 EFFECTIVE Figure I.Sound .2 .4 .6 .8 1.0 MACH 1.2 NUMBER PROPELLER as level a function of number Mach numbers. of propeller blades and [] [] 8-BLADE 2-BLADE GEARED FIXED-PITCH FIXED-PITCH ADVISORY FOR AERONAUTICS DIRECT-DRIVE NATIONAL COMMITTEE MAXIMUM SPEED CRUSING SPEED TAKE-OFF DISTANCE _///////////////////////////////////_. I I I I _///////_ I I 0 2 4 RELATIVE performance eight-blade and 6 8 1.0 PERFORMANCE of an geared airplane fixed-pitch equipped propellers. with 1°2 Figure direct 2.- Relative drive two-blade 1D D D t_CONTROLLABLE I_ CONTROLLABLE []FIXED-PITCH MAXIMUM SPEED CRUISING SPEED TAKE-OFF DISTANCE 8-BLADE 2-BLADE 2-BLADE GEARED DIRECT-DRIVE DIRECT-DRIVE l_i_i__i_i_i_i_ii_i_i__i_i_ _////////////////J_/_ ..... '"'-'""" _ _/////////////////////_ I I ..... --"""---_i_i_i_ I I I I I 0 .2 .4 .6 .8 1.0 1.2 RELATIVE Figure 3.Relative performance PERFORMANCE of an airplane equipped with controllableand fixed-pitch propellers. C NATIONAL COMMITTEE ADVISOI:IY FOg AERONAUTICS Figure 4.- Two-blade direct-drive and eight-blade geared propeller configurations. %.76 ". D_AG CLEAN-UP \ . lowing air flow through the ductin_ system. wing._g the strca_m[_ne nose that was added on the NACA open-nose cowling and opening the cowling exit so that the engine cooling air could flow added 18. the clean-up of a typical airp]_e will be described first and _. and cockpit enclosure remains._.. Remov_..atcr some selected detail cases taken from thres air_]_nes w.6 percent..ndition._accessory compartment added 3 percent. A_d_ng the unfaired carburetor scoop added 3. t_il surf_co_'. During the war the most important work of the Langley full-scaletunnel section was the drag cle:m-up of military alrplane_. The airplane with all the dtems added theft pert'_!n to the power plant installation is sho%_a in figure 2. installing the proJecting exhaust stacks and opening up the hole throu_tb which they projected increased the drag 3.. We use the dr_g of this b_sic condlt_on . openin@_ the outlet at the back of th. Item by item th_ airplane is retur_od to its s_n:vlco condition and the drag is evaluated at each steo. TO DETAIl. Uiluecessary drag wrLs found to result from (]) projection of various items outside c._i. In order to show the relative importance of these _tems and the locations _n which they are likely to occur and. Points The first step of the clean-up is the at which it-is suspected that there analyzed carefully so that The airplane is then put examinati(:_n of the airplane. all openings have been closed.!ling the Intercooler scoop. Instc.as reference and call it 100 percent. increased . and all external leaks have been sealed. Herbert Wilson. it has be_n summarized in two repo_s (referencesl and 2).f a smc_oth basic contoul. DESIGN Jr. is unnecessary drag are s_ud_..6 percent of the drag of the basic c. In order that this work could be made available for general use. changes to reduce the drag can in '_ faired and sealed condition in which all protrusions have been removed or c_refu[ly f.red. For ex_ole.6 percent. also to acquaint you w!th the procedure that was used during the invostS_:. Only the basic combination of streamline body._d and be designed. These reports include the results obtained with 23 airplanes and point out the advantsces to be obtained from careful detail design. 4W_ " \. and a).ations.(9) roughness of surfaces (3) _nintentional leakage of _ir through the airplane struct1_e :rod (4) the us_ of large quantities of excess air for various cooling functions. fid_ure 1 shows such an airplane in the faired and sea]ed condition.77 PERFORMANCE GAINS By BY ATTENTION A.il] be g_vun. °._ry.erial adde_ 4. In order to illustrate the principles of clea:n-up in more detail a few addition'-_l items from other alrplan_s are shown._ke it usefD1 we have increased mostly large._ Just sho_a'_ wottld increase its speed by about ! re:lie pe:_"hc..sions._ts drag nearly appear with! at what has happened to the clean airplomo order to _.cving sepals fromthe gaps on the cowling flaps incre_Lsed the •rag _._.t_rms of incrc:_cnts of dra_ coefficient.s th_ incr(ments e.6percent _nd the d_'ag of the rc.8 percent._ st_.ly cle_n and in b rir_:In_z it to a usable configuration unnecess_a_Y dr. the _Ealn woulcl be sli_ht]y loss than 1/. and leakage totals 19. It is partlc_larly impc._g was added alon_ w_th tn._ccn altog(thcr. by addle4 items that by 65 Dercent_ particularly the_.8 percent.._D. In all these cas_. opening the case and ]._rd.n t._ mi!eK in sp.. instal lin_ the radio _. wh. \ . th._ __n from the same 4ra_ increment -_euld be 2 miles per hour .:t.c-por-h(. Look In roughness. _nd adding the machine guns amd blast tubes added 1.. l_ge outlet increase_ the drag 10._d :)n an q i. We st.hour class or about an 8 percent increase this amo_u_t'bd to_._Tp!'uqein the 200._'ted with anu airplane in fi_re 1 th_. an_. opening the s_al around the edges of the land_ngge-a_'doors increased the drag by 1. A_!diticnal tests and careful 8nalysis sho_od that the drag of tlle pc>_'er p]_t items could be r_duced to .t_r. necessary to _ut t].: in _. A&d_n_< the sandsurfaced slip-_roof walkaway added 4.ane _n _ In figure 3 _t S s _._rted .8 percent. however.2 percent. thus s_ving nearly On this a.uto gave h 1._ drag ass(:clated with the necessary functi<ms.:. Yet._ percent.rtant to note th. w._. For yo_r inforv_tien it is pointed out that a red:_ctlon of 0.t was exceotion_.2 percent. '] .neral those _tems have dr_%gs of nnly a few percent each. they add up to an impressive total. This second group of items which include protr_._..78 the drag 6.6.per-hour clause.n an airolano in th_ over h00 miles per hour cL.6 percent and installing th_ oil-cooler system consistin_ of an external scoo_ an_ an urmecessaril_.2 percent. is not necess. These items total 45.kage items 36 percent of which is per hour could be the drag reduced to 9. Next.ss.mlle.2 percent.tw a_-_'P -]-an. u_moss and le.9 percent_ of the basic condition._n that the re_.lrplane in the 3"25-milo-per-. the rest of the items service condition are a_ded._selves dc_ not All of this drag..c :_!rp].0002 in the drag coefficient of the BOD-m]]. (.ink eJecto-_' c. mile _ per hour..regiven in .6 percsnt of the dram of the basic condition. The 1_n/_aired i_rotrucion _d excess air leakage c%used t.stable landing gears.n airplane._l. The _ cover was then extendc.. with the blunt shorttailed fairing which extended shown by bothnose th.0099. The inst:_llation could be greatly reduced b. i .r_.md_ng sear retracted is shown in figure 7.m that would cool the stacks. It was provlde_.]'reducing.s to chose the holes and provide adequate and well deslbmed cooling air outlet area elsewhere._ng _d tail reduced the dr_g coefficient [_hown 0... Also in another case.f of the figure the original prod_ctlon installation f<.r th.'med_V _si_tisfactory cooling of the engine for the climb condition.e cAw.%rwas used. Tile low<_r h_. Sealing _ f_]Ir:Ing this ex_haust opening decreased the drag by 0.d to cover the well.'79 The upper half of figure 4 shows _-n engine exhaust stack of the ._ sketch.ls alr]_lano. Half wheel well was left uncovered and the roug_ hole ther<_ caused flow and hi{h drag._ing Just behind the cylinder baffles in order to r'. On the lower half of the figure is shown an e:<haust-stack installation that does not protrude._e_ tyo_.'wed t_. by only 0. and directing the air re al_-a.:. because of the high pressure difference between the inside and outside. The remedy in this case _s.in the cowl'_ng_.? speed by aPl_rox_mately 13 miles per hour.0041 for the cowlings of th_ two entwines of the airplane. Addi_..he t mint_n_._s n_t "when the i/8-inch shows of th_ poor _nt..:.!re but se_iL-ed.o.sto_er.._on wo_ld increase the airp!an. Engilm operating tests made both with the original stacks installed and with individual jet type exhaust stacks showed that the c. Opening '_oles. single case in which a fixed l_ding gr. This reduced the _rag. the cowling near tube incre_sed the Althou_'_h we 'ordinarily expect high performance ai_Tlanes t.:_mbined reduction in drag and increase in thrust of the cleaner insta!la_. In fi_re 5 an airp!_ne is she_n in which holes were cut througlt tl. either before or behind the engine is especially ba@.[_Is_nstallatien to hmve high drag._ir was all. • _e-l. but w.0010. the longer fair. dr_g of this alrYlow to the dlschargcd Some makeshift methods of prov_dln_ e_zine cool]nr% ha_e been fotund to cause e_'_cessive drag...ssive amount of ._ have retry.0008. figure 6 shows o. cutting a hole through the nose to get clea_-ance for a mach_ne-gunb!ast drag 0.0003._which was one of two _-:n_. f!_w out of this opening at a high angle to th9 free strew._L_. The fomu drag is not large. The h_]os failed to provide the desired improvem_nt in cool:h_g and ful]-_c::_le--tunncl tests showed that the flow disturbances caused by the holes increased the drag coefficient 0. gap around the edge was sealed_ the drag was . but an exc_. An example of the more normal case with the ]. mple that an airplane could be cle_ulcd up to tl_e extent of reducing its drag coefficient by 15 percent. LSA30 .-ies inversely as the cube root of the drag ccoffi cient_ this would give a speed increase of app. In conclusion. I want to point out two things further.-oxim__tely 5 percent. A St_mary of Drag Results from Recent Tests of Army and _avy Airpl_cs. Also. Dearborn. whereas the personal aircraft des__gner is likely to be intei'ested also in the horsepower required and fuel economy at the cruisin_i{ speed. The military airplanes fle_. in contrast to the much ]ower speed 8_d wing loading of . There are great diffe_"ences in the characteristics and purpose of military and personal airplanes._ and Silverste!n. Now considering the horsepower required _'or cruisin_ which varies directly in proportion to the drag coefficient. . ROy H.sortable. L_ingley .s the spe_ds anal g_nera]. the _mportance of good detail design will increase.cale-Tu_e! N&CA ACR No. Suppose for ex_. 19kO. C. 19!_5.8o decres.an airpl_m_e that cruises on 150 horsepower.: Fuil_. NACA ACR.sedanother 0. let us consider what the app3ication of these principles to personal aircraft mes]_s. H.rat speeds of 395 to -_'_90 miles per hour and had w_ng loadlngs of from about 25 pot]n_s per square foot up. _..0012. Based on the res_its obtained with military airplanes this seems red.. Abe: Drac Analysis of S!ng3eEngine Military _irplunes Tested in the NACA F_'ll-Scale Wind Tunnel. F_rst that the gains that will result from caref_l attention to detail design are likely to be greater t/tan the differences _n drag bet_een airplanes of different cles_ basic confictu'atlons. desi_n affect h!g_l-speed airplanes.5 horsepower for ._FERENCES I. Oct. The incl'ements have been applied only to the top speed. it is seen _hat a 15 percent reduction in the cruising drag coefficient would be a 15 percent reduction in cl_uising horsepower or 22. perfol_nance of personal aircr_ft :_dwmce. Langc. Since the speed va_.in the detail. Now th6_t you have seen how refinements . necessity for preventing random leakage of air This emphr<sizes the throu{zh the :&irp].ersonal aircraft. 2.0009 or a total ef 0.ane. --4 °.3i6 3.. NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS . airplane items with power-plant-installation added...0 v. Figure I. drag Figure 2..6 PLANT. 45.6 6.0_n .v 3..- Clean-up airplane in the faired and sealed condition....- Clean-up .-_ (iii: \.6 . - Exhaust-stack installations.8 :35..2 POSSIBLE REDUCTION .2 POWER PLANT -45. # L 4 NATIONAL / 4 ADVISO_Y FOg AERONAUTICS COMNITTEE Figure 4.-.8\ 18.7 Figure 3.4 4. .6 LEAKAGE AND ROUGHNESS--19.2 "'' 10.o 5.= 1..o 3.6 = _= "" _...- Clean-up airplane in the service condition. 64.6 3. .- Fixed-landing-gear fairing.. _-...-" NATIONAL COMMITTEE ADVISORY FOR AERONAUTICS Figure 6. / .© co /j j_ Figure 5..//" .. _.. "°'"_ J / .. ?/ i / ... ..Cowling with holes cut through behind baffles.. © le--_ NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS Figure 7.- Retracted-landing-gear fairing. . ki AII_6_A_'_ LOAIX_ .61 y. thathas been learned is applicable to light air%enos. i]._l dimensions _rithin which the velocity chnnges take place. the lo_. and the gust most likely intensity. Fi. dsta.he gusts but it requires further a h_o-_ledge of the .._llfall scatter regardless _nto-the of the c_. It within the plot therefore indicates the size of assoclated with a_.. on throe airplanes rs_ging in size fz_m the _mall Aoronca light airpl_e to the XB-15 bomber.1 airplane._wts having the intensity shown. _[D . plott. much of :._nes.cture n o _. f_'om measurements of acceleration and __irspeed.ely the velocity (_radients and the spati. The.-of ._ez'o obtained._n _:m. especially those in the nonacz'obat!c _Id transport catego!"ies.mube'..ega_lless of the category of the aiu_l_e.y given value of the gust i \. For many _ircraft.%o the persons.nt d:_sig_u requirements are be. It mey thus be concluded that the gust st:_-_.:m_st stru_t_._'e.STATISTICS OF ATMOSPHERIC GhoSTS Ill _RELATI01_ TO By Charles TH)] FERSONAL A_RTL_ C. The .tz_ structure.8P_ SOME I_A_ ON TEE STRUCTUre.ro gove_'_e& by the s_me principles that apply to o_he-.:h!ch the c_trrc.dit is given in terms of _Ing chord lengths. The effective gust intensity Ue is plotted _. of a ftuadamental nature._e H.m_ep_-ttern size of the airplane. the investiG_tions have boon.ds imposed by _ists _ro critical in the desi_.ga_nst the gradient distan. od _.m_ctur_ . : can be se(_n that the limits of the the data _. Shufflebar_:er C: The flight loads on personal airpl_nes s. Each point on the plot represents the most probable gradlont distance of a l_a-ge n.-n of the main _'._ed is an applicable to the _mall a_rplane as to the larger tz'_nsport t:FPe_':. however.gure I su_uartzes the resttlts of some of our investigations on .7 types of aiz-pl. Our research on loads therefore is.:a _-eq_irements on gust loads are the same and the question arises as to whether these requiz_ements are applicable to personnel alrplaues. . the C_A desi_.gradient distance in the distance _rlthin "_hich the velocity or intensity of the gust changes from zero to a maxim_. in general. applicable . neve_theless. Whi!c the Leboratcry's invesblgations of gust loads have not been directed primarily toward establish: ug what those loads are for po]_sonal airplanes. so thaT.. The design of airplanes to withst_nd _Dist loads requires not only a h_owledge of the intensities of [.nc_.g_st st. that the dgsi_ gust intensity of BOfeet per second now incorporated in the desi._th the turbtulent regions. The meas_'ementsw_th the _(C-35were taken within tbuudei._ length _ud it applies specifically to average transport conditions for which the path ratio R or ratio of miles flo_.83 Figure 2 further su1_iates the almplicability of our fundamental gust inv--_gardens to the smell aii_L_nes.eater than the correspondin_ 6nlst intensity shown._n requirements is reasonable as awplied to the personal airplane category'. The two frequency dietributions sho_u_ wore taken fr:_m reference i and represent the frequency distr'_burdens _. For example.renc.istributions of the effective _st velocity° These f__-equency distributions are s_m._ne :_d by measur_mentswith an Aeronca C-2 airplane.to the gro_und._s between the airplanes on which the measurementswere ma(leand between the meteorological conditions associated _.s. The frequ_ncy distributions. and ctuc statistical stu_lies have centered largely _bout the problem of the transport airplane. The curve in figure 3 taken from reference 2 showsth_ result of absolute frequency determinations for transport ai_q?lanes. The absolute f!-equeucy or the numberof gusts of a glvon intensity that will be oncountgred within a given period of operation dependsupon the ratio of the miles flo_ in ro_h air to the total operating miles florin. The absolute frequency is represented by the inverse qu_ar_ity distance Me in t_.5 _!rp].'_n seen.e 3. as c.rms of operating miles roqui_'ed to encounter a ghostof a Given intensity. There is someevidence to show._. are be n_arly identical notwithst_a%dingthe considerable diff. The cu_re sho'_nis given in to_ns of unit chc. The m_asuremcnts_ith th(. which determine the ttu'bulent _-eglons of the atmospherethrou4_hwhich the various classss of alr_lanes will fly.plya plot of the relative frequency of occurrence of gusts _ose intensities lie within certain !im_ts.storms at all altitudes b_tween about 5000 an_% 35..ete_ncd by moas_romerits with a Loc!d_eod XC-'_. .. Statistical det. C-2 cn the othe.'-.el_ninations of the frequency distribution of !i_ustintensity az_e often of value and sho_." hmud were taken in turbulence caused by s_arlng of the.10. _n_d clos__.o) _lll have an intensity equ_-l to or gl. a relative f_. This evidence is shown in fi_ro 3. The dist:_zcc required to encountcr one _ust of a given intensity for an_vvalue of the cho_:dlength and for any w_lue of the path ratio can be dete_nined from the curve and by __pplication of the formula shownon fi_n_.s on the operating conditions. however.000 feet._ here are the relatiw frequency _.quen_zy lO-_ meansth:_t onu cut oY of 1CO0_Asts encountered (on the averag._ in rough air to total operating miles is 0. The actual choice of design values of gust intensity dopend. This question involves the statistics of operating practices. n the following marn_er: An arbitrary obsolescence operating time for a typical transport airplane :_s selected. the relative frequency distribution of gusts. requires some modification from the volue applicable to average transport conditions.ll or moderato chan._ty o._at lower altitudes than is common to average tr.m in the inset table were determ_Ined _. when incolq_orated in the formula and -_ith }. For example. ho_eve. airplane. In this corzecti_-n it mi@ht be noted by reference to the fl_. In conclusion.lO to the personal ai__plam. Some of the . for example.i. we would be tn a position to say that the design gust intensity of 30 feet per secon_ _as probably too large to be correct for the personol-airplane category.'.-hen operating altitudes _mre in the neighborhood of 3000 to _000 feet indicate that for such conditions a path ratio of 0. _e path ratio R.84 The operating miles M an& the operating times M_r sho_. we can solve for the operating llfe and compare it w!th that originally chosen for the transport ail-glane.t airplane ra_her then substa_utlally smaller. that the operating life for the personal aiz_lene worked cut t_ be substantially lazger than that of the transpo_. respectively.lO corresponding to average transport operating conditions. if we choose a person_A air-.. If we had feared. something less than half of the operating life chosen for the transport machine..earlier transport data taken _..n O. to or exceeding 30 feet per second within its operating life.e operations. These n1_bers. _J_e s_ppose that the personal airplane will._ taken at a value corresponding to a gust intensity of 30 feet per second in effect stipulate that this value of the gust intensity which corresponds to the 1 .and the choz_l length and operating speed of the airplnue were taken t. When this w_lu_ of path ratio is used in comp_ing the onerating time for the persona/. on the whole. taken as O.%nsport operations. a ch(%nge in gust intensity of only about 6 percent _rlll result in a change of the flight path by a factor of 2.s 19. plane snd permit it to enco_.mit load factor will be equalled or exceeded a given number of times within the selected operating llfe. The valuechosen was _°O. be operated at some. that smo._s of operations is found. the value of about 8900 ho_. __ne only si_wnificance we attach to this result is that the _L rL•i• operating lives of the personal and transport airplanes work out in a proper relative order.ght-pe.000 hours . the design gust veloc. feet and 200 miles per hour.20 is more nesa-ly applicable th_. Now. and the epera[:ing times for the tre_ncport and personal airplane. The path ratio _s.ering gust st_cture. ccnsi_._ter the same number ef gusts equ_.th c_rve. of course._os in the _st_intonsity result in fairly large ch_%ugos in the oper-_ting flight path._' 30 feet per second appears to be of the right order for personal .. Philip: Frequency of Occurrence of AtmospDe_'ic Gusts and of Related Loads . A_ny Air Forces. L2:I21. Reginald Statistical NACA . and CIu_._msport airplanes.0n Airpls_ne Structures.et._R No.85 as weI1 as for t1-... ust-Load Criteria. B. L6Cll. l_ichard V. m_d Donely. RFL_ERF_CES I Rhode. : An Application of Data in the Development of C. 194_. 1946o 2. especially for It the the should also be mentioned design loads for a nonacrobatic types. T D. Bur. _d Reise. Aero. _land. NACA ARE No. . that the qust load will probably be number of airplanes. . Ue.....(D GUST INTENSITY.. I..P.Gradient distance of gusts determined 2O from measuremen[s with three airplanes.) FiGure 2..random C-2 Relative-frequency distribution for samples selected at XC-:"...._nc:_ from turbulence surveys with Lockheed airplanes.. FPS 50- AERONCA XC-35 XB-15 SPAN 56 55 150 CHORD 4 9.. XC-35 IN ! I(_ _ _\ THUNDERSTORMS NATIONAL CONNITTEE ADVISOI_Y FOg AEgONAUTICS h Id 'J.. 20 30 -20 -IO O GUST INTENSITY...._ '<_nd Aer'.. ..S. -30 -1 .o cd .-_. I..Q .2 18.. .(9 b...... IO (Ue--'F. .. FiGure I..7 0 E3 z_ i 0 GRAD.O-- -4 RELATIVE FREQUENCY _ AERONCA IN WIND SHEAR-_ G-. r . o (1) F_ DISTANCE.1 1. Aw_rage airline operating conditions..Miles per gust equal to or greater than selected values for airplanes with mean geometric chord of 1 foot._.3x106 E 950 PERSONAL I0 TRANSPORT 150 200 2.(1) b. 0.I-_ FOR ONE GUST =.Ue ! iO-I o io 2'0 INTENSITY.ooo _M=.2 0. HR. 4o Ue" FPS GUST Figure 3. NATIONAL COMMITTEE ADVISORY FOg AERONAUTICS . Mo "MILES PER GUST 105 FLIGHT TIMES AIRPLANE c FT 8 12 V MPH FOR Ue =-30 FPS M W( _/V R MI. Some work has been done in this field and the rat. the necessary stability derivatives were obtained from model tests. In fact.il has been in the specification of t}:D con*r<.._. Owing to 1-mitationc of time we will confin_ our remarks to what has been done and what is available for the determination of the loads on horizontal snd vertical ta. it for computing the design load on the hori_.of . Other comDarisens (see reference 2) which have been made for ai_l_%nos of various sizes also indicate good agreement so _at with regard to the horizontal tail load. it will be _een that the agreement of the loads on both the wing and tail is exceptionally good.onshad to do both with the detenlin_tion of the load mag4_. The method requires a knowledge of certain stability derivatives which may either Be obtained from the available wlnd-t_nnel results or may be computed by ordinary engineering procedtu'es. such a method has already been incorporated into the design requirements for military machines. it may be stated that a method exists which is not only rational but which gives good agreement _mder usual conditions. a fundamental part in aTplying. much has been accomplished to rationali_e the problem and to relate the tail loads to the geometry and stability characteristics of the machine. Wi±h regard to the determination of the magnitude of the horizontal tail load. Although a suitable method exists.s for the _-51 airpla_le in a p__l-out m_le at about 315 miles per hotu_. These ±n_e_ti%:abi. In this connection. a convenient method has been developed (reference i) by which it is possible to compute the horizontal-tail load and the wing load for any given elevator deflection. Figure 1 shows the agreement which is obtained betw(_en the calc_lated and experimental tail load._AFPir_'.itudes and the load di_t_ibution._son During the wsr a m. _ic calculated res_llts are based on the actuc-1 clow_I_?r motion used .in fllgjlt.AIVD EOPITD_f_L TAIL S_RFACJ_S By Henry A.l motion to be used._nber of ]oad investigation_ wer_ undertsd:en _gncse recults are applicable to the personal airpl_e.8_ _.'O LOADS ON VERTIC_JJ . It r_ght be mentioned that in this case.-.cntal t_. Pe'_.s.ll surfaces for airpl:uaes in the acrobatic category. imer. f In the up_er left ha_zl corner of fi. for this particular machine the ultimate design ].#hich it is moved forward rapidly in t_me to check the mo2euver at the specified load factor.'.necus rudder k.ve ex_mples apply to a larg_ flying bo_t flying at 200 miles per he'. and of 33° at the endpoint. (See references 4 and 5. (See reference 3. The top right hand portion of flg_e 3 shows a rudder operation of 1° made in roson.vPe of airplane.kick mad__.'ontal tail. of only 15c at the midpoint.the correct tail load can be calc_lated. For design.) The results of these investigations ha_-e made possible the specification cf conservative stick motions _hich _re within the physical capabilities of the pilot end which can be made to tie __nwith the specified load factor._ure 3. as w_.87 control movement and the forces which a pilot is capable of exerting on the stick have been invest_gate_._ith the hori._._hat happens to the vertical tail loa.. Directly below we see that the load on the vertical tail at flrct increases by l_O0 pounds and then. that providing the rudde_ motion and aerodynamic characteristics ar_ kno_. steady value of about 700 pounds._el! the calculated _nd oxpel. Thus.e effort has also been directed tow. tall load problem and two pertinent reports have been _sstle_._.um v. the load reaches a maz_.'Iv_tives f the airplane.s the case _. Therefore. f_nally _£ter several o_cillations.tive control motion is the one in which the pilot pulls the stic_ beck as rapidly v-ud as far as is poss._neuver any vertical tail s:_rface could be broken providing the m_euvcr were continued sufficiently long. following . While the illustrai tf.._nce- .oa_ for which the tail was des.u-. the most censerv-_.) As in the case of the horizontal tail.c and st_bility de_.i _ith different types of rudder motion. it seems evident. a _Teasonablo agre_zment was obtained bet_-een theory au_ expe_'iment. the results of s_me computations are p:.. In order to illustrate . attaining v. _zehave _m instemt_._a_ rationaliESng the vertical. It h_s been though% that in the fishtaillng _. It is seen that as in the case of the horizontal tail. th_ results are qualitatively applicable to any t. lone. o F_gure _ sho_'_ how _._ble. as the _. The dottsd line shows the computed load using the experimental rudder motion._ck of lO. the methods "_4qich haw: been developed dopeud on beth the geometr_.gnedwot_J[dbe reached by sn _nstantaueous deflection of about 19 ° at the initial condition.'esented n figure 3.. at 2_0 miles per hour with the P-hOK air. tal vertical te_il loa_Is agree for a ruddei.-_lueof l_O0 pounds. So_.Irpl_me s_ings toward the equillbri_n pc_s_tlon. we have reached the point whore the specification of a reasonable elev_tor motion is possible. In fact. is more difficult than the srccification of an elevator motion because there is no well established lateral acceleration to which the rudder motion sara be linked and also because resonant effects are likely to be more pronounced on the vertical tail load then in the case of the horizontal tail load. we can say tha. the subject will not be covered at thi_ time.t with regard to tail loads adequate methods are available for computing the loads which are __mposed for a given control motion. _. it would ttltimately reach 2800 pot-ads. By comparing the two cases. . it is possible to build up the tail load to larger values than are shown by imposing combinations of the two types of rudder motion. however. It can al_o be seen that the tail load approaches a limit s_ does not increase indefinitely as might be supposed.¢ith regard to control motions. it is seen that the type of rudder motion has much to do with the vertical tail load. With this particular maneuver. whereas.o8 • _th the airplane's natu_ral period. Summarizing then. Large vortical "tail loads also arise from other mo_euvers sttch _. The lower rig_ht hand curve shows the tall load compute& for the above r_dder oscillation. A2_thou_h we have not reached a simil_r point for the rudder motion. Yt is seen that in this case the vertical tail load.sa ___ll_ng pull-out( see reference 5) as well as from direct _Adder action. reaches approx_uately 2700 po_u_ds per decree of deflection'. Since this subject is covered in one of our recent reports (see reference 5) and since rudder motion is not a main factor in rolling pull-outs. _Aftor about t_ cycles. The logical specification of a rudder motion. statistical data of the types of motions used by pilots are being obtained. if carried on ludefinitely. or_ly about 9° rudder m_vement would cause the tail load to reach the ultimate design value. 5. Henry A. NACA _B No. Boshar.J_enCalctulated aud Measured Loads on Wing and Horlzoatal Tail in Pull-Up M_euvers. L4HI4. John. NACA A_ Jan. : Comparison bet_. NACA TN No. 1_5. Philip: Consideration of Dynamic Loads on the Vertical Tall by the Theory of Flat Yawing M_ueuvers. L4E31. Cilruth. Robert R. 19hg.89 I_E. .: Analysis of Vertical-Tail Loads in Rolling Pull-Out Maneuvers. Matheny.: Deriwztion of Charts for Deto_minSng the Horizontal Tail Load Variation _r!th . Cloyce E. 2. 1944. 1944. Beeler. 1065. 1943. NACA CB No. and Davis. De Re: Maxlm_mRates of Control Motion Obzained from Ground Tests. Pearson._j Elevator Motion. 4. 3._ENCES i. NACA ARR Nc LSHll. 0 and measured vertical-tail motion. 1.6 2. 0 . LB.Comparison of calculated and measured tail loads accelerations for an XP-51 airplane in a dive pull-out._ o 4ACGEL. load on a P-40K airplane due to rudder . ALv.4 TIME. An. ALt. g 2- O: 5xlO 3 2TAIL LOAD.---'i ! I I EXPERIMENTAL TAIL LOAD. . 2 SEC.. LB. Figure 2. I I --MEASURED ICALCULATED 0 -I- I I I 0 I TIME..2 FOg AERONAUTICS ! ! 1. . DEC and -20 -4O 0 t I / "k. RUDDER DEFLECTION. 3 Figure i.8 SEG of calculated .Comparison . -----CALCULATED 400800- //_ NATIONAL CONNITTEE ADVISOI_Y o-400_S / -800 4 . ASr. o o_ RUDDER DEFLECTION. ACr. COMMITTEE . (o) RUDDER KICK --4000 SEC (b) FISHTAIL 0-I 000-2000 6 0 . DEC RUDDER DEFLECTION. 12 . . 24 .0 I __1 I I I I I FINAL I I I I I I I I TAIL 2000I000- LOAD. A8 r. 16 . NATIONAL ADVISOI_Y FOR AEI_ONAUTICS /" L. LB/DEG TAIL LOAD.. DEC (/ VALUE 2-0 t 1.- Calculated vertical tail loads for a flyinE boat in rudder kick and fishtail maneuvers. 20 . LB/DEG MAXIMUM 40002000O-2000. 2 3 4 5 Co TIME. Figure 3. _ . 1 ? _B . t _ - k.9o t. about l_ or 16 pound.the a_tainm_nt of adequate take-off pe_Tormance. Parkinson DESIGN 7_y John INTRODUCTION During the wsa'.gn. research on seapl_les _n the Langley tanks was mainly concerned with the large multi_ngh_e flying boats used for patrol and transport missions._cause they operate at higher Froude n_mbers._uuamic problems -minor or were information. and _._i+_the m!nim_n of penalties _n aerodynamic Different combinations of wing loe. water r-_sistan_e is a r.'_ field. s. power loading.less of size or confi_n. Little direct experienoe was gained _.ith the amph. the data and conclusions _n the Wartime Reports from th_ tenks _re a?pl_cable to the porsonal-airplan. hi_her speeds in . h_drody_mumic stability end controllability.ud arr_augement alter the rel atlve importanc.v _m_mportant criterion in hull dosi<_n as cemparud to stability and spray. in fact. Attention was centered on hydrodynamic stability and sp_-ay problems associated with the overload!n_ long-range (reference of existing types. s_d thus accentuate other hydrod0u_amic problems° For the same wing and power •loadings. it appeared th2_t related to their design and oper_. that is.nd the Hullos Hercules 1)._s._ing._obi_ms. being adequately Basically. small soapl&nes tend to >_ve poorer take-off performsnco than large seapl_mes b.91 SF_2LANE AND AMPHIBIAN B._er loadings. but they must all be properly consld_red to pro_uce the most successful desi. and on the design of very large transports llke the Martin Mars s.. the designer of . and provide additional background for post-war developments._lativel. / CONV_TION Hydrodynamic AL R75LLS Resistance R9lstive imp_ertance._ of the w_rious hydrodynamic requirements.Jght. prolong the take-off run. however. In this ].tion were solved on th_ bas_s of existing s_ller _eaplanes and hy___od. it tends to limit th_ useful load. seaworthiness perfor_Lnce.ration .:rater-based airplanes is confronted with the same general requirements regard. With high po_er loading.s _or horsepower.With low po. _. Effect of length-Tbc. The relatively low wing loadlnzs usually _sociat_d with personal-o_er airpl_s. Research along th_.. the improvementin take... A typical affect of increasing length-beam ratio whil2 holding the length-beam product coz}stant on the total rcsist_unce durin_z t_ce-off is sho_min figt_e 1.udbeamconstant.-off performance would of course b_ correspondingly greater. halo a favorable effect by re_ucin_ th_ load on the rater and the take-off speeds.u#_ by tank inw_.._ased the re.sist_ce at bo_h th_ humpand hig_hplaning speeds._ of !ine_s has been conduct_od a wind-tunnel a. Consequenbly. not only toward reduction in the drag coefficients of the hulls or floats. (unpublished) of a rebated series of hulls with len_Eth.'nncsis usua!f!y infc._ll size without tuudn_ impa_rm.8 by _ncreaslng length an.'fo_" --manc_ muot be diructed. With a hig_her power loading and smaller take-off thrust.. designed to have comparable hydrodynamic characteristics but _zmnllcr size as the icngth-be._or to th.s.s____!Ta_t_io.nt the hydrodynamic qual_ties.in performanc. but also toward rc_duction in size by the use of more hy_odFnamically efficient bottoms. the effects of the s_ze of the hull must be e]_aainated as far as passible when Investigatizk_ the effects of length-b_amratio alone on take-off perf0zm_nce. on the other han_.slng be&_as indicated decr_._mratio is increased.ments in pe'.92 relation to their dimensions.-b_mratios of from 6 to 12'.z.The water resistance at the h_ soec! and at hlgh planing speeds is largely affected in opposite directions by the hull size for a given gross w_icht. w_th a resulting &'_creasein take-off time of 12 percent.) Increasing the length-besm ratio from D. lmprow._ resist_uc_ at various length-besm r_tios is to hold tZqeproduct of the length 8.The aerodynamic performance of scapl. ._in length-boLunratio offers the alte_n_atc posL_ibility of reducing drag by reducing h_.1decre_. (Data_is taken from referenced 2. Onemethod of holding slza constant vhen comparing th. Aerod_naamic Drag Si_gniflcance. Effect of !ength-beamratlo.2 to 7._stlg:tt_on .t of the coE_espondinglandplanes b_cause the bulk of the seaplane is g%_eater provide spray c!ear_nce to for propc_llcrs and acrodynamic surfac_s. This proccdto:e is analogous to holding the wing area constant when comparing th(_ drag at various aspect ratios.-off ._with increas_._.The Smprovem_nt tak_. and the __ntegrated effects of d_sign p. the. Those models provide a close simulation of actual seaplane operatic.. and thus becomes a useful criterion for hydrodynamic stability. r_ducos the drag by some 20 percent. 6. The hydrodynamic data for the series are not complete but indicate the length beam ratio lR hull to be as good as or better than the length-beam r_tlo 6 hull. 5._gley 300 MPH 7oby 10-foot tunnel. The over-all trend on the aerodynamic porfo_m_Ico. A number of reports on inw_stigations of dynamic models arc available (references 4.hydrodynamic performance. These characteristics result For a given hull. and important hull variables. but _ corresponding increase in verticaltall area.All conventional unstable on take-off if trimmed below the f'/ seaplane hulls lowor. Hydrod_n_nic Longitudinal Stability Dynamic models. Increasing the length .beam ratio from the conventional 6 to t_o novel 12.porpoising are limit or above the upper-porpoising limit. trim of gravity and the elevator in a range of stable positions of the center of gravity for take off analogous to th_ aerodynamic operating range of center-_f-g_ravlty positions. depends on the position of the c_ntur moment._ndicate a possible rnduction in horizontal-tail area.vity " range are shown in figure 3 in which the maximum amplitudes of porpoising encountered during accelerated take-offs arc plotted against the fore-and-aft position of the center of gravity for . which cover most of the Take-off stability. They demonstrate clearly to the design engineer the relative importance of various hydrodynamic criteria. The stable centerof-gravity range can bs obtained in a straightforward manner in the t_nk or in flight tests.Research on the porpoising of seaplanes has been carried out in the tsnks w_th powerc(l dynamically similar models. The coefficients are all based on the wing area of the same hypothetical seaplane and ropros®nt the absolute changes that wo1_ld be obtained by interchanging the hulls on the assumed design. when the leneth-b_am ratio is increased. The slopes of the pitching-s_nd yawingmoment curves (Cm_ and Cn_) . appears to be favorable.%rameters on these criteria. Typical data establishing the hydrodynamic center-of-grs. and their characteristics cr_n be directly correlated with pilot's experience (rof_r_r_c_ 3). while mainteining comparable .£3 The resulting hull forms are shown in figure 2 along with thepertinent aerodynamic characteristLcs as obtained from tests in thQ L t. 7). howow_r. " The skipping motions In tr_ and vertical position.Several flying boats form of hydrodynamic instability and amphibians have during landing (and in some _nstances during take-off) called "skipping. the hydrodynamic center-of-gravlty range is a useful criterion for this important design parameter. the hydrodynamic stable range at this elevator setting is from 30 percent to 37 percent of the mean aerodynamic chomd. is not of the favored by Because of th_ large effect of the fore-and-aft position of the step or. of to the elevator. and the seaplane is often thrown clear of the water below flying speed (reference Research skipping and that 8). Assuming a maximum allowable amplitude of 2 °. Unless allthe aerodynamic and hydrodynamic moments can bG closely estimated.9_ one elevator position. ventilation deep of the afterbody area sufficient ventilation behind the step. the hydrodynamic moments about the center of gravity. As the center of gravity is moved forward past the point where lower-limit porpoising is encountered. If the hydrodyn:_mic range does not correspond to the aerodynamic operating range. respectively. a dynamic mod_l having insofficient At low landings whore landing wer_ smooth the aftorbody depth trims with landing stability. Landing a-serious _omponents of the airplane are closely had stability. likewise. the ampl. particularly for small seaplanes. however. Consequently. Figure 4 from reference data were obtained from of step for adequate where the maln stop little amplitude to attain The 9 illustrates the conclusions stated. in which all the moment-producing simulated. the amplitude increases progressJvely to a dangerous value. touched first. if the center of gravity is moved aft past the point where upperlimit porpoising is encountered._tude also builds up progressively. the proper location of the step is best determined by dynamic model tests in the tank. dynam:Ic models in the t_nk has shown that with is primarily thecause of a function of the landing tr_m the instability is insufficient directly in this and speed. it may be shifted fore and aftby moving the step. either by an adequately step or by auxiliary ventilation ducts. _ea. the At higher in trim.i _ movement elevator pilots. Extreme use extend the stable range. trims . the step is located largely on the basis of the desired forward hydrodynamic limit. In practice. This stable range may be shifted forward or aft by upward or downward _ .. must be provided stable landing characteristics (reference 9). are rapid and not under control of the pilot. A_u increase _n depth of stop iron 5. against le:.able landings can be r:ade only at low triii_ and high.w depth of step.2 _ercont _' of the beam eliminated the skZpping for all practical otu'Doses. d. This exper_ence and a .hs that the l_-ncth of the forebody has a pronounced effect on the c!eannoss of rt. Spray Charactc_ristics R e!a_i._u% at low water speeds over the be'#.c"o-I__z_ti_.oss-load r coefficient and the forebody-lcngth-beam ratio as follows: = -. sati_factory Low values low-spsod of sp'_'ay into the propellers._uuin6 at taxying s)_eds: also that chan6es in bes_ alone httve a relatively minor effect. even though the choruses represent l_Go vc_riations in. s_. th_ relative _rit of other variations _n step design ir_fluencing the effective depth. con be evaluated in terms of an important.07.-It has been co_:.nd forebody. The chief hydrodynsaLic _ena].drodynamic quality.95 touched first. Similarly.05 to 0. 0. The corresponding data for the amplitudes similar trends with landing tri_: and of vort'. The relation between depth of step and landing stability pro_rides a useful criterion _'or the _minim_:. that is.=orre!ation of the ]. and k.l_. shallow ste. depth of step required. l_undlng speeds (reference lO).= where the gross of feet of forebody from bow load._ to 7.mon experience in the t_.u_own spray characteristics of multienglne flying boats (reference ll) have led to an empirical relationship bet_een the g._n sh_.i_. length sea water_ 64 pounds per cubic foot to step. usually .beam loading.on of h_'d_ro_%_ic loadin_ c. pounds AO _ I" !'_'h (1) density b Lf and the be_.ien__cions._d 'to the flaps. feet k is a coefficient that varies approximatol5 linearly with severity of th_ spray thro. or the hydrodjno_uic effects of fixed step fairings to reduce drag. the model skipped violently at all trims up to the s+_ll. easily recognized h_.ty for a full step fairing is the s_e as for _. 09 _hich limits over a vide rsnge of hull proportions. &o/Lf2b i_.96 characteristics values.Observations of the spray of heavily loaded flying beats and their tank cl_ly _he necessity of powered propellers for tank . The validity of the products L_b and Lf2b as fundamontal from tank hydrodynamic parameters has been further est_!Ished tests of the ser_es of hulls referred to previously. 0.'plcal d_t_ establishing. I_ spoed ran_. while high to O.11. Both set_ of boundaries . there are two r_gions corresponding to light spray which do_s l_ttle damage and main spray which _s of most concern. also results _n constant trim tracks and r_sistance ov_r wide variations in length-beam ratio. _. usua31y correspond the over-a_l usefulness of to _n_desirab!_ spray the seaplar_. but. A constant value of the product L2b_ where L is the length from bow to sternpost._kept the same.tests to adequately reproduce the full-size spray pstterns and to d_termine fully ths effectiveness of spray-control devices._ _s a function of load over which spray was observed in the propellers of a twin-engine dynnmic model are shorn in figure 9. the rotating _vopoller blades and slipstream greatly increase the height and volume of the spray at taxying speeds. The minimum _ro_s load increases approxlmately linearly with upward movement of the propeller position (reference 13). the spray coefficient k has proven to be a useful criterion for the selection of forebody proportions and d%mensions of a multlengLne c cnf iguratl on. An investigation of the effects of various pa_rametors on the spray in propellers of a f_ying boat has shown that the application of power reduces the minimum gros_ load at which spray strikes the outboard propellers by some 20 percent. and reduce the amount striking the tail surface at higher speeds (reference 12). It is interesting to note that from equation (i): &O wLf_b Hence same k for and the severity of the spray remain if the more ratio or less the different length-beam ratios. For normal configurations. In general. There have been exceptions to the rifle. Effects undesirable modnls show of powered propellers. _n seneral. _ll combinedwith spraycontrol devices as required to obtain tolerable spray characterist_s. Tests of models have indicated that the planing-tail hull has muchlower resistance than a conventional hull (references 14. The speed range and minimumload for spray in the propellers or for spray striking. The d_gree of conservatism in this respect dependsof course on the type of seaplane. TEENACA PLANING-TAIL Ht_L The results of the type of research O_conventional hulls described have lead to the development in the tanks of an unconventional form called the planing-tail hull. they are Gasily determined either in th_ powered-modeltests or full size. S_a__ control.ished results of investigations with powered dynamic models have indicated (ii_: . which maybe of interest in the personal-airplane field. and 16). unpub]. forebody spray strips which have transverse discontlnultie_ to break up the spray instead of smoothly faired transverse sections to build uo a pressure peak at thechlne are best._ forebody-aftcrbody length rhtio also has a favorable effect on the spray..97 converge with decrease in gross load and cometogether at the loads for which the propellers are c]. and thus provldecomparable data for correlation purposes. hence it is useful for high power loadings or for reduction of hull volume. The most efficient aerodynamic design can probably be obtained by use of a small overloaded. _@_ile not direct measures of the severity. Increasing thetrimby increasing the sternpost angle or th.Various meansof keeping spray out of the propellers and off the flaps have be_n evaluated in the tanks and have been applied full size. A sound and not over-optimistlc estimate of the maximum gross .other componentsof _he a_rplane. In addition. This hull has a deep pointed step faired in plan form and a long aft_rbody extending back to the tail surfaces. Similar data maybe obtained for spray striking the flaps and tail surfaces. One of tl_e _r_ost ffective e forms of spray strips investigated w_s a vertical plate extending dc_nwardfrom the forebody chines. In general. In addition._eight in the _reliminary design sta_e is essential to the optimumdesign of the hull lines (reference 1). 15. The best method of insuring satisfactory spray and seaworthiness for a seaplane is to maintain the proper relationships between the weight to be carried and the dimensions of the hull.oar. are useful criteria for the effectiveness of spray-control devices. as comnared to the usual value of _.5. depth and location of the step. landings same order A orofi!e of the new hull form and a comparison of' its resistance with that of a conventional hull are shown in figure 6.6 for the conventional hu]. Further inLorovei_nts in the personal-airplane field may be achieved by more attention to the form and size of hull. load-resistance ratio at the hump speed of 6.stance of the planing-tail by virtue of the large afterbody clearance. balancing of the longitudinal moments. and soray characteristics of a new desi_u-may be estiH_te_ approxil. the plardng-taJl hull has _. of CONCLUSION The hydrodyr_unic resistance. hull. In the comparison. feet pounds (:/_b> per second . resistance coefficient {wb"_) R CV V water speed water resistance. feet (wb_" J % W sea pounds pez cubic foot b CR of hull.98 it to have superior dynamic stability on rake-offs and and _ind-tunnel tests have shown its drag to be of the as for conventional hulls. stability. SYMBOLS One gross-load gross weight beam load. the fee'. end relations of the loads and proportions.mtely from the generalized data and experience found in the NACA reports.l. At speeds near take-off.. coefficient speed. the high. host that has been recorded. of coefficient pounds water. is of the order 50 percent of that of the conventional hull. Bnt. dynamic foot-pounds foot-poun_s IM\ _q-_# M q S C PreSs'm'e.-mament coefficient. pitching moment. variation of ys_ing-mcm. feet per second 2 ratio coefficient with angl_ of L_ C length-beam variation of pitch_ng-moment attaak.nlt coeffici3nt N b y_wing mom.acceleration of gravity. foot-pounds wing span_._. pounds per square foot wing ar_e. square f_-et me_n aerodynamic chord of wins. feet C yuw_ nM-mom.nt coefficient with a_gLe of yaw C pitching. feet w CDml a D drag_ pounds . : NACA i0...s Jn Fo_m: of Hull on Lcngitudinal Stability. Ll.. 3KO9.i.Size Flying Boats.s Det. 5J24_ S. NACA AI_{ No.-tt_ Start.C&T.saJ-.ns bh_ For ob_..rkinson_ John B.gltvdin.. Be!i.f !. of Sc. Locke.Rang_ Flying B. 7. M}dol Depths Nor_zn S. _TACA V-Stop _d wit_ No. 4. p_. F!ying-_oat Model Equippe_ D_signcd to R_duco the Air No. James M.n. C!':_.L9. Robert F. 3E13. 1943. john of a Po_. and DyrmmLic Two Land. 923._. I -.: An Analysis <. IIACA _IR He.Tnvos_Jg4_+ion of L. Bonson.I.ity Bo_t with a 30 _ Ste_. Havens. 194['. i_\CA ARR. W.r_. mud. John B. : F2fect of Po_re_l rropo.. Parkinson.btho_s _Js_& for 5h. E.: of a Flying of Transverse or_ th.mrod B.: Tz_J_ TEsss of a with Several _. Joe W. NACA "_I_L'o. F.rud:w_mic cf a Dyn_" No. ]$43. Nov. 8. and Olscn_ Roland E.ity h. Jr.: The L<mgitudi-ual Szabilitj of Flying Boa%s _'._llfi_1944. !_nd. Norm_.bu Skipping cf Some Full.'. II . N_CA ARE No. 1945._ of a Long._i-Stabi].rao_:cristics.i1_g B._ A<. 5. P'_.o.n._i. I<ACA APR. a T_rgu 2. IiACA AIIR ll. on the Skipping Char_c.ENCES 1. Ho_ar([: Efi'_ct of L_ng_._E 9.. I_ John No.. Normom S.i00 IRE.: The Longit. 4E!4_ 1944._. Charlie C.-_t._nin¢i by T_s'ts c£ M_d_Is in the _'_. ap!an<_s. Trusc.!i_rs Characteristics and the Porpoising Stability Mod. Garrison.. ani I/_nd._ .:zt_ with Spccial Roforonc:_ tc PcrpGislng an_ Skipping. : Long-Range Flying The DGsi_ of the OpS_mLmt IIull for Boat. 1944._ Critcrions_for the Dimonsi. J&mes M. B_nson.. NACA 1946. L5CO9b. 3J23. of .: Dosi. 6. Notes 1943.. NACA RB Parkinson. Oison_ Rci_td.] of Fairing Drag of the _ain Stop. .._minoi by T-sts of Models in t h_ NAC% T_a_k..u_ S. B.h-Boam Ra_:io on Resistance _md Spray of Throe Models of Flying-B<at Hulls. Tho I_nding [_tabi]._llna! Stability of Flying Boa_s as Det. 3.EffecL el" V_riatior. and Zock. tcri_tics No. Park_nsc._l of a Long-Range Flying Boa l. : 3127. : Piloting of F/_. Effect of Varying Le_ngth. ar_i Gross Load. Robert. : Tank Tests to Determine the Effect of Vaa..NACA Model 11'7. .. Robert C. McKann. John B. Propellers . Tank 15. and No. 1943. and Hay. Dawson. Kennebh Tests with Planing.lOl 12. Dawscn. John R. II .ying Design Paramo_ers of Planing-Tall Hulls. Elizabeth S. Elizabeth S. _lan-Form Taper of Afterbody. Dec. : Tank Tests of a 1/5 Full-Size Dynam'ca!ly Similar Model of the Arnkv0A-9 /_. John R. _L_CA %N No.. Width. 1096_1946. Angle of Afterbody Keel. John R.: Prelimin_ry Hulls. and Ha_'.: Tank Tests to Determine the Effect of Varying Design Parameters of Planlng-Tail Hulls. NACA %_ 16. L. Roland E. 13. Motor-Driven 1941. Jolm R. NACA TN No. and Wadlin. 1946. Parkinson. _19. 1062. and Olson. 1946. Dawson. : The Effects of Various Parmuet_rs on the Load at Which Spray Enters Propellers of a Flying Boat. Dawson. a_d Walter.. Walter. NACA ARR the l_. ll01.. I ._.Effect of Varying Depth of Stop. Robert C. Length of Afterbod_v Chine..-Tail Seaplan_ No.ph:Lbian with _ACA ARR. P. . D " 12. 90 F. CDmin .0075 G'm a .8__( / it'\ • ----_L --52--_ b i " _' .by I0-foot 300-mph wind tunnel.0009 .4OxlO 6 Figure 2..0012 NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS .-_ LB_ 36.000- cfl 0_ L 24.0060 .- Effect of increasing length-beam ratio on the aeroin the Langley dynamic characteristics of hulls as determined 7.0067 .Typical effect of increasing length-beam ratio with constant length-beam product on take-off resistance.0 F F--__ I I I I I I 30 60 SPEED.000_ 7.S. © o. 120 150 Figure l.N.000- / _ TAKE .0044 .RESISTANCE.0062 Cn$ ."3.0051 .00135 R. .- Typical effectof contact trim and depth of step on amplitude of skipping during landing. NAT IONAL ADVISORY MAX.A.2 I . I I 4 7. DEG.. 40 Figure 3.Typical effectof lonEitudina/position of the center of gravity on amplitude of porpoising during take-off with constant elevator Setting.o o3 .5 4 _ I 0 I 8 AT CONTACT.C._-/ AMPLITUDE o. 32 PERCENT 56 M. AMPLITUDE. ALLOWABLE 2.G. AFTERBODY HORIZONTAL KEEL CONNITTEE FOg AERONAUTICS STALL I I I DEPTH OF STEP.o DF-G. 6_ o_ LOWER LIMIT POR POI SING UPPER LIMIT P ORPOISIN G _ STABLE RANGE • ° . TRIM Figure 4. DEG. I 24 POSITION 28 OF G. AMPLI TUDE.. % BEAM 5.el MAX. . Gross load coefficient c%__. . SPEED 4 ' COEFFICIENT .24LOAD-RESISTANCE 4°6 RATIO AT HUMP COEFFICIENT. Figure 6. LB.0 LOAD. Cv 6 .GROSS I00 . PROPELLERS _ CLEAR 60- 40 i I i 0 8 SPEED. IN MAIN SPRAY PROPELLERS IN LIGHT SPRAY PROPELLERS 8O _x. _ CONVENTIONAL 0 :. 24 Figure 5. RESISTANCE .Typical effect of gross weight on the range of speeds for spray in the propellers ol a model of a twinengine flying boat._. I. 16 ERS.o. OR NATIONAL COMMITTEE ADVISORY FOIl AERONAUTICS / ..Typical comparison of resistance at best trim of conventional and planing-tail hulls. 102 C POWEI_ PLANTS . irpl_ue flo_. Vs. n of _ six-c:_lindcr Franklin air-cooled f°lat engine of 130 horse. Rubert '_NS Kennedy The radial f.tring_nt ruqulremonts of the proposad target airplane.f air-cooled engine has had such cx_.[_s.s a_d baffles designed and constructed _d_r the d!r_ction of the.0.representative V-type a_._0d_l 33 e..r-cooled _ngines resorts& in r_f'erencos 2 and 3 and problems of l_ou_d-ccoled -"-.craft. t..inlet for cooling alr _as provl<ed directly below the propeller axis.ing test bed for the co_.. NACA. Th_. deslgned to !ncroas9 the available cooling pressure differential and to decrease the a i_--flow and pressure-drop requ!rem_ts. inv_st!_tions fez._-ryapplication w!del$ used and has co_li&mt_ations have of optimum cowlings that tha NACA ra<lial engine cowling has been undergone intens iv_ devcloP_ _. installations are treated In references 4 an. This project was conducted in the fall of 1941 for ti_a Army Air Forces. '?'. ls shown In f[g_u._-_vo nil:._on did net provlde suff.nal ai±.-.na _t Lan_le_¢ F!oI&. A Fi_etwings _.._!:i_. _nd. This Installat.oht airplane class.. powered with :_ s!x-cylindor flat a!r-cool¢_eng.he free stream wuc obtained b2 loca_!ng the inlet as far as possible from the axle of rotat'icn._na.power for a r_:.io5 AERODYYAMICS By OF _-aD'_ER. d_alt w!th the development of cowllng for an a_rplan_a of the l:i.Ic_._tal pre_m_re in excess of that of _.lcicnt cooling capaclty to meet the more _'.-PLANT INSTALLATI F.-an:. The cool:iug _:[r imsscd downward: across tha unbaffled cylinders and was discharged at the sides and bottom of the cowling. 2.dlocontrolled target airplane. Th_ fundanent_l principles of cowling d_sl_. The application of these principles will be illustrated by a discussion of a research reported in reference l_ wi_. 0ti_er eng!ne found favor in person_! aircraft_ but uho use for these Las been less gcnet'_l.. Cooling alr was supplied to the bop of the ong:[_ t_r<_u_ nose _nle.& 5. . Materi_l Co_maud. The revised Installation._rm . b_ compemy personnel was used as a fly.nt. in flgure 1. Similar are Th_ investigation chosen to illustrate these prlnr:l[_)o_ _as the develop_uent of a power-plant installatlc. by the Fleetw!ngs Co_a_y and th_ NACA wor_-:[ns in close cooperation. A s_ngle ._a!rplana before modification !s show_.al_ methods of analysis are tLe san_ for all englno conflgt'a_at.ts on each si&_ of tha propeller shaft. the eroerlence ac_ulred in _evelcping cowlings far militury aircraft ma_ be applied to the improvelaent of cowlings for p_rs.. _ cub.. lJmiU_t!on.:. the uxtorn_! '__ _"' " a'r sbrc:.sb{L..Lng colt].larich.tip]._:_tl' =.)r the _ngino...lv..e& bj -the inlet_ h._.!octJ:.:.:o_: duo to the absence of b<-un&ar_ ].. C passv. _n(t coL_"_(U':-n< l.d.._a be _ ... !".n.u climb con_lit ton. Ch.s:'.:.._._Gtions with &_...%s mast b: inc..otJ.aycrs :rev! Imb w.el}ace.d in th.t shown was placed :.y_ t_LO &i._ cJruum_ta:.or .3s__]b]..:_z" o£ I'_.. _ 'at higA _:.'oa! _u:J.%kes..u is f_vo±'able to stabilii G' (f ° i'low at i_:w-hJ. az_& an inlet v¢. _..it throu&._::_"]Y_' distri]sution of flow _mong a _m].['li._.lat..j.'.nl of gui_.._c._._ of .t plan. :L_cl introduce& it into a _nurous prc.. thereby grca..& b._"_.n_.i.].C[ !.nabi_ the {].:a.<.5ysbo..irc_ct!2tmd.(_[: J_._ d_t:T.w_!]._ nccos::.n._bor C<.:' was a]most ._1_Id bO r_ilCh _LEI:.'..._.ions c:.:koE:r.cai.] wou.-..pj...:._r of' ._"_ ..:oity rat.j..'.:_L:.'.mi_f a.gs x_or_.-:sp".inJ.ssos.:as so io_.::. or piunm_l ci_.h_.-rs tv r_s_r:_cb tkc £Jow 1:0 that llSU._s is G.gi'._.:. oxJ.c[ty ratios ar. _.{<e on _h._91!'_. .:.3 was i'<:ul[d pr.)sc_m..._c_s'.!.i(:.%Liv_ly small co].]:."_ aa ofi'Jc].-.1 in:ta]..za oha_fber w-'.d.jf OI!_..::_w is a pow...i..tiTat.et proper a dif£m.£at.te& bG n. :bi_is c}l-':._/ hq<z].. " .loci% r::tlo ol 0.ut:.'_blo pr::.:<..c_. iv.$h l_su wc.pz'o?s] _n of ere.t " Ic.t op_._st.'c]Z.'. fit%& to _no cylin&.. In this ?N. : '-" _ of ".:d. Air from the p:h...mb<::c l-:. "L 2.] location chose. Baff].ct L.' van.iesir_..ntr:)._'ry for stab:!. [:._t_ly aft of i.iLxl./7.J !&'r_:o cL .:._LTt).avo].d.uvla.rab!.s pr:_vJdod wh".cated _].abio intal..evi:_5 _iu d.p o.:. This location except in a p].:' } h:. a_i& thu [_r cov. tic r.-:_ :)Y c:_nw::.Z ]:'in_: u. .L!f_dlsUG_:_Ub _:..o.udLev_'.....L._. d'.:_ntir:_J.:c-b:i(...r l.]. .d ]. t_.C..'aploye&: but for most app].fzh perf'oz:maI. 'id.L_na..y i_tac_c :pressure.a. or bc.f<:.'.[Ly of be'_+. -I:L.:pl:[ c:_ted and ineff ].._ ]:<.f.:c.r_s_]t:.?[.d_ an_ for pil.'-._d from the b_f!'l_s flowed to a /'!:<..II pas._:rful deterrent to ±'J.l_.r:3.I_ _he point.nd_..tt.:l • _ _ of dynamic pressure to static press[_.-.9 is essuntial ]..'t.::...oi. Th_ favorablo pressure (:..! :.}_ a r_]..._< alt:a& o [' the . f!o_ in the 2H.bc coo]. be av.ropo. of the 0.."Or " ] "' _ " t.ausc of wind:_hio?.i.css craft b_c'._ l._.t sys t.._zorced f uoward acrus_ I.3 the more L%8 active as part.tLJd ba o]:..._c}i_l'_ c._hox'¢ spizmel. usuz'.t _:lJ._tv.id offe..mit_& n__fo._Or ['dlaZl th.?.nd to ol::minate all .ighor in\or vcL._ could b._'{_'.'.er'.:].m...s o:.. The air dischar£:. Ti.fficioncy.:: in the ta_r$.h.e flox. _x:i.h_ m:.o'.104 benefiting * O _:hcreby from _. c[usi._d fl..h_L_.u .bl.ci.bi..n "bb.: cy!in&ers_ in v_vorso o±" tn.ces.i in all dir.i_[ condition._h_._ary. k_ .ic.: _'oioci (y ..__ )i' this principl.i_j_::.bLaln.i.rectb:n of the or!.._nt _¢hich occuz's :n :_(:c_:!__.vLJ.:I _x..roi.::..:_... f flow !r_.:It[.ssure loss.Ller disc..nC t h_ {' a_rc_af_ s i.lJ_.._ino..l. Immcdl.nbl)y i._.. although quite chc_rt h_causo of spac. wi.['the _ng'no.. I:l ]_.tin_g <ver [.ct I " " <: v._'!_ :i:' ::_ c..c:_'.r.¢t:!.'U]J._ /n].rcc i l'cr &. wa:_ .y_ the greatest posE!._ssi'av:. aiU_ou.'e :.a$cr :._ i':x.. as _. ].._ad.._:rW.c Lively still f_xrther roducocl bile v_looJty of 'bLo :[nc.l& Lc cisbribute.{as _.:t.:'. of/_.a(_ Y)r in+¢rnal diffusion -_ '_ _ calm. Apl?.. .ch..: spz_.. Station-by-station a powerful aid in the analysis preliminary of the internal-_'low of power-plant systom is of design installations. A_oidance of critical-flow regions in 1ucatin_ duct exits eliminates another common source of dra_. designed for adeq. For deflections up to lO ° or 15 ° to the s_face approaching the limit of flap effectiveness. The exhaust stacks as a protection air by the hot which discharged below tho cowlln_ were against fire and to prevent heating of stacks. or to the hot-air stoves. incrgasing thereby the presstu-e a_ailable for cooling. &t extensions which are quite frequently encountered._cy in climb results in gross overcoo!in6 in hi_h-speed level flight with consequent excess cooling dl_ag. The socond function of the hinged flap is to create a local depression of the external static presst_e as it is extended. The use of a fixed opening. Dual carburetors were fed from shroudsd the cooling a manifold k : open at the front to the ram pressure entering the plenum ch_nber and connected in the rear to the exhaust shrouds which could be used as carburetor. Excessive extensions. Free swingin_ doors fitted with an override control permit connection of the carburetors either to the full ramming intake. the first of which is to adjust the area of the exit oioening such that tile minimtu_ needed cooling-alr flow is obtained.. A hinged flap performs two functions. r_y so increase the external drag that the loss of flight speed docreases the aw_ilable dyr_mlc pressure more than the decrease in static pressure :_btainod at the exit. . In such a case. Careful s_aping of the external cowling lines to eliminate pressure peaks or local hi6h velocities insures a smooth external air flow.air preheaters. Figure 3 shows this installation with the addition of the exhaust and intake system. which in a body of adequate fineness ratio gives rlso to little drag. however. reference 6 shows that'an exit to give adequate cooling for a lO0-mile-per-hour sea-]ewl climb is responsible for a cooling-drag power in level fli@ht at 220 miles per hour of seven times that necss_ary. Reference 7 reports tests of a representative flap installation which caused external drag equal in _otuAt to all of the rest of the drag of the airplane. Backfire control was obtained through slammin6 of the freeswinging doors and relief through a backfire door at the bottom the cowling. this can be accmnplished without serious increase in drac. For example. This particular installation is of added interest because of the ease with which other features fitted into a design primarily dictated by cooling considerations.lO5 are far preferable. extending the flaps _lows the airplane do_ and overheats the engine. .ai..'ht l test &at& el' ti_o installation describe& and halo an acc_a-acy much better than cou_Id bo obtained by osti_'lato al(... As th:_ coolinc _J.data _n ref_:r. co_ ling l'an_'.y be omp loy6d.aom_try is fa_.r._.lIra_:sm:o _:_bthu dueL.zo6 and is equally imnortant in the evaluation of exi_ting configurations.9_Jp.d fli_:.suro .r exit..a exit prossu__e. !s conservative 5 to n:ogloct this boost.-..>.n<_. d:_scussod in r_foz'ences lO and Ii oi'f_r the most <. Wh_ro the g.:]. i3 described .n the pressure after &ischarg_ returns to that of th. particularly of tk. shows the t the &re C c_%r_c tcr._ronce _. The in:Lot ar. correcting.ndition of the air is el-deed stop by step from tim free strg_ to _he pol_.huory .al]at:Lcn u/frierdiscussion.c _oa__l_ _. tits cal. Th_ calcula:_ions of figur_ 4 wor_ mad_ from corr. difference and ._ding to 1/3 horsepower at th<_ llO-milo-p.s obtaln. fc. coz._if!oant . Th. _xtornml t_iu static calc_latod.cl _mder favorable oonditions_ lacking exact informe.s [c t _s well _¢iti'_antolerable limias._ctors _.'.fbha inul.roin.o tt't_.._. and quantity of discharge._vpu which n_y be r_quirod to taxi f<r extoi%&ed periods ab l:_:irly h__gh power need some m_-:ana for au_nting the coolin(_ a:[r fl..b:gwh_.rgeable to the internal flow. The resultin_ information pets.r _-_sas t_._nco 9._i' oxh:._ou6h tbe prop:9!l:. Th_ internal brag in the case illustr&te(i is rou.tions such _._suro lossgs are applied to calo_. DuctlnP.. exhaust cj. if n:accssary._ly 1 po_i.s that dcsorib::d ._am.in detail in ref. the {_esire& in3_eh-volooity raai. which is th. is i!lustr_tod in condensed l'<_r_l Ln ficure 4. ..3 first steps in a pr_ll_in_r_ ins ba.. and th. temper_ttu'e._l pres[_u'co withzn th.u_ :-:(. exit.:r internal &re6.mss flow and the chang_ Ln velocity from t]_t of flip. In ttnsy_muGtz'i<:a! ins_alit. &ucbJ._sical cond._oolJng--Passa_> _ P l'_.nfz.ce. Internal _ao_ ls calc_ulat_d as bb:_ product of _.ll_tion analysis are setting the p.ustc:j._m_in._charg6 velocity _:_&vbo necessary exit area is then obhaino& dir_ctl._ coddling air quantlty an_. This low valu. Such an analy_i_ be_segt on fi_{_:t tesas _-.b_:J.:ct._d.lysis tom system is [livido< into elements in each of.nd obtaining the coolin_E air flow and press_ure drop requlrod by the enginr_.o c<.o_.or:_bl_.lon.gr-hour flight _3po.iis_:b-_r. In this uypa of a_m._'._l's i_5 given in reference 12'._._ at low fob-ward _p@cds. . lo_so_ c_z_ l)e calc__/aficd from ._se_}.at to thai r_.'r a boost in total pra..changes of conditibns of the air flow._a is caloulate_[ from th.'.cul&tion cf t_m drag force and power ch. the i:i.h%.-.l... _¢hioh it is possible to state the si_. Pusher aircraft. ti¢_n J._ f.lut._to Ll.c pr_cs_n-<_ of .ltions of pressure..t . The method [email protected]. From the d Tnam/.e.ce_ _tr. and speed of the frG_ sbrc_m_ _.. _dlioh however ere Usually sr:_ll.._f discharge. f.. A t. TPe from oh9 velocity between tot. and.. references 14 and 15 on coollng-duct _xit flaps. by mea_sof which engine-c_. .i07 Other papers of value or interest to the installation englneer include referenco 13 on cooling fin design. which gives the NACA cooling-correlation theory.. reference 16.oling air requirements are related to operating conditions. references 17 and 18 which present practical application of the t_eory to multicylinder engine installations. _. Robert D.. : and Cooling Tests of a Fleetwincs Model 33 AirT>!. for Radial _..'crcos. 1939. 8. John @.ing Power. Her!_ert. NACA ARR. Ed_rard: The Prop_ller Charactoristico of a Twin. 2. L4H15. 662..zo8 REF_L_]NCE S Cuwling i. Wind-T_znnel Inw:atigation NACA AER No. Cz'_rnecki. NACA ARR No. 3.tion {d" tl. Pressureof Duct Components. Arn_y Air l. No. Hem_Y. of Rear 6 K..z.o R<unger V-770-8 Engine Installation for bhe Edo XJSE-I Airplane.>!.: Ducts. Cooling-Air-Flow and Pepper.. Abe: Review of NACA Investigations Engine Co. A Me. 1943.thod for tho Pow_r-Plant Inst.. 7._ and. Nov. R. l. nd Nelson. and Harring_ton.'.alla- 9. N_%CA ARR No. 12. NAC_. and Keith. 1943. Jr. Silverstein. L4F26.ation cf the Cowling of the Bell XP-7I' AirpLanO in the '_£'.C. J.2_[al-Flow 1944. Underslung Fuselage W.ropellorResearch Tunnel. : Full-Scale WJr_l-Tunnsl Invsstigatlon of Forward Unders!tmg Cooling-Air Ducts. Nelson. NACA ACIt No. : Dssign Los. Konnedy F.:tT. NACA MR. 1943. with Propellers of NACA lb-Series Airfoil Sections. Gus_.in£_. George W. az_d Den_rd. on Fans for ii. 1945.. NACA 2J_R No..A. El!erbrock. March 1942. L4120. Nichols. 4. Nichols.TED No. . Czarnecki. W. lJ. I0. TM . Mark R. Jr. Arvid L. VI C. 1944.. NACA 044.'.: Design of Cooling Systems for Aircraft tions.L4_ -].. NACA [. MuLterperl._ No. july Wil. _n_ Design of Jet Pump_. James G. Rubert.: Design Air-Cooled Ea%gines._ Charactoristics 1944. Herman H. of Power-Pian_ Inst_llations. J. 312]. N'.%_<: :!n Flight.. McHugh.zcnle. NACA ARR..Engino ancl Airplan_ Mod__J1 Equipped with NACA Ds-TYpo Cowling'.. Fans and Fl_2gel._ Coolin. II . C_-org_ S._:bb.. NACA MI_.. 19)i4. Ma_. Jan.. and Wilson. s.: A_ Inv<. 194]. 1944. Mark R. John R. Bur.i_u._f!. R.A. A. A_ro.Aerodyna_¢ics . _CA of N. K. an& !(nopf. Cowlings Rep. ?_c:. : Anl InvestJg_. LSI12b.u: High-Altltud. 5. Jr. Arm_ _ Air F<_rcos. . 1944./ . Blake W. No. Arnold E.941.: Tests of the XSB2D-! Engine Installation in tlm 16-Foot High-. Charact<_ristics Processes in A!r-Coo!cd No. NACA S. Alhq No. Biermann. Charl_:s He : Cooling of a Pratt. Jr. Herman H. I'$ACAR_p.: Cow]i_l(_ _ lap_-...Jy R-28o0 Entwine In_-jtalled Cox¢ling. Cerson_ Blak_ W. L4F06. 18.. Ellerbrock.. V . 16. 14. 120. 1941. Katzoff. Aero. in an I_\CA D s Shorl.Nose Hig_h. 1_44.Cowling and Ductin_'. Stickle._ Wh[tn. : Hig)__-Altl_u&e Cooling. John I. David. George W. Lhll!d.I:Ll¢t-Vslocity NACA ACR No.. and McL:. Cr_-glur.. NACA R_p. Pinkcl. March '> /- < ._!lan. ?'26. Jr.. Jr.. • ¢) 17. 16%CA . Irvun_ and Pressure Available for Cooi]ng with Rep. Benjsadn: Heat Transfvr Engine Cylinders. 1938. and Corson.. Nai_m_n. No. Bur.Speed Tulnnel. 6L. ]. NACA MR. 19. Biermann.Io9 13. : The Desi_ of Fins for Air-Cooled Cylinders.. . r r . C&OLI_ AIR INLET Figure 2. as developed .- Fleetwings model 33 trainer before airplane power-plant arrangement. installing target- ._ _._U i' ' Figure I._ AIR ' _ i<.i " '.Cooling COOliNG NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS for use on arrangement target for cowling airplane. VELOCITY.lJ_i-i_6 ii.1-.2-°38 l 2o4.2..AIR F FLAP-_._ E) HAUST (iii COOLING AIR CARBURETORS CARD.Induction and exhaust airplane system cowling. 2070 Iblsq ft p.AIR (COLD) Figure 3. STATIC PRESSURE. '4 -4 I -22 I 0 I 0 Internal-flow analysis.1 E-----_ZL--I_I . . fps 165 Ap Figure 4. CARD.-. 2041 Ib/sq ft V.. TOTAL PRESSURE.- ° I1 I 5 _I4 / 5 / o864 2084 Jr2°6. 9-1 i. AIR (HOT) NATIONAL ADVISORY COMMITTEE FORAERONAUTICS arrangement in target- 4 5 0 H. _dL'.nd .twc_n 32 ° F rmd 40 ° F. and on the carburetor parts.. _Je w. due _d Throttling ice forms whenever the effoc t:! v._ll first _on_'sid.zed b.m_'<.L m_.Jning bet_ecn private ar.ct ice.d by vsnturi suction. character:ist Los.tor ventur:! throu_h a partly closed throttle _s oncu_:h to cond.: quickly become block(_d under these imr_.r b_low "32 F :impinge ° and freeze on cold surfaces at the a_.n. c_rb._'_ct_c_ w_ll form whenever visible water droplets "n ai.l .i[ i.y rcp:.ately warned by the presence _e priw._ 1. Im_.d transport differc.i_r thcsc present a s_mi]._t wil.to pilot will.d <. cyliodors. a carburetor which meters the _.rrt_ that in 1949 there wore over 195 forced landings of pe_r..r _nt_ko. Inc_[de the duct.ng difference in perfo1_n_ice is attributabL_ partly to inherent differences "n icing characteristlcs of the two inductlon systems ar.cm_] .._r_lP_ne _nducbion systems _. SYS_L}. as compared with no forced l. fovm_tion w_!l increase with a_rspc. let u. Analysis indicates that this strik'.Jzast_nces.m_ in this :_y_tem: imp._!c_! li'gl_t airplane _nducticn system in which the fuel f_c.t oreva]ent abov._l more iuclement weather. p_'cportions the fuel spray _.5 F..:.r. The basic elements of this sjst_m com:..s (v:. _kny air filter or screen in the cold _-_ir :nl.:r<_ on the a:!rcr_ft.{rature drop tc the .d partly to differences _n skill and tr_.]L b. but the pilot is of Lice c!scwb._rwlse. evas_vc action un.. _-._" :('c ._r oJ' adenitrine fuel to bh_ system. throttling ice._dlncp of tr_=isport aircraft for the same cause despibe their oper_tion tbrou.ar probl(._ons.xp_rimonts _:how that this phenom_non can occur in clear a_-' but ls ].r]:_c: a cold alr intake.'I-3 }!unt_r The need for better ice protection for light n.pes of ice can f'o. usu_lly take circ._ Most conventional induction elements but differ slgnlficantly systems in tho contatn th._s clearly reem_has]. The rate of "moact ic. First.s/c mann._ s%me b_'.n't.ns_: water and freeze _t at or beyond th_ throttl_s..t<_mp.n_foi d for a conveying the fuel-air mixture to the engine._p_ratur_b.m in generally serious oth.'_<_}_minc th..':c!t'cr._ratur_--_ well ° below 0 ° F.._d and w_. Thz'_tt]i_.era ty.L ] ICING OF ENGI}_ By Wi!ison I_DUCTiON H.i..nd throttles the air flow.L. Dry snow and sleet the case of the f_iter but are not condit._.qft <> due to _his fo_m of ic!ng. nces in icing pilots.u_:_-. Three distinct ts. end fubl-evaooration ice. Analys]_ _. but can be e__pectsd at tcmp._tcr contcnt_ wi.-'mitcd to !nlot air te._xpansion of moist air through th. throttling process is most important.t the even distribution of mixture to the cylinders. Fuel-evaporation ice forms whenever the air :rod adjacent :e_ta. The. together with any ingested rain. but at part throttle. because it contributes at all temperatures to the more prevalent and serious fuel-evaporation icing process..L surfaces are cooled below freezing due to the _vaporatlon of fuel.es in nhe tt._erin.:_se in fuel-air ratio or by fuel tcmpcratur. :_ince a finite t_me ]s required to evaporate fuel. power may continue to be lost . It has be. however.d. ic_ may form on the fuel nozzl. Large amounts of impact K._ndfu.:n learned that fuel-evaporation icing _s acgr'rcat_'d by mor_ volatile fuels but is not si@n!. a loss of alr flow will be detected as a loss of engine rpm . or eny particular geographic area. rbulent region beyond the t/_rottle in a m_mer that both increasers and localizes the re2rJgorat_on effect. When this occurs additional wat_r condenses from th_ air and freezes.m _nd is not necessarily associated with bad _eather. little effect on air flow at part throttle because the system is th<. or clouds. freeze. but they may also seriously affect fuel metering. it is to be r. in th9 system we ar_ cons_.t can be seen that induction system icing is a year-around probL.l flow. the throttle. fuel Lmpinges directly on the throttle plate s_nd eddJ.z. until forced landing but cruisiz_ is n_cessary.ci6ntly to obstruct fu_. The throe forms of icing will all produce a loss in engine alr flow at a g_ven throttle satting.iii c_. %n alert pilot who does not suspect what is happening will have a t(-_nd_mcy to a_ivance the throttle to regain powc_ ".._. However. This process can continue intermittently until the throttles stick or until full throttle is attained.xpected that :in a rapidly mowing air s bream serious cooling would develop only well beyond the fuel spray._ itself suff_.-_am of the point of _iniz::_l pressure are cook. At very low air flow or high water concentrations. obstruct small air f During cruising flight at constant altitude._%dairplane spc. or ups_.u fo_ :n any system in _h_ch nt_tal surfaces do_istr. In a well designed system operating at full throttle this is actually the case. This explains why inexperiunced p-_iots are so easily fooled by induction-syet_'m icing.:_._devaporation Ice before or after the carburetor will cause.n handling amounts flow and only a fraction of the design air flow. d below freezing. Thur '.ficently affected by ]ncr_. Our ezm_erim:'nts Jndic_t_ that in many induction systems fuel-evaporation ice occrrJ in clear air at inlet temperatures as high as i00 ° F. of ice on the throttle edges will quickly may freeze the throttles. :/_twould be app]._nts _. It w_]. and if the engine _s still developing enough power to provide s_sffic. In ess. however.duced. it c_u become iced and cooled until cor.!._s for the pSlot to exercise Judgment at the proper time to d_2-ice the system and prevent furth_-_r :icing by m.. to merit any consideration as When sm engine is idling or being operated on the grot_14 for long periods at low power.:: will be consid_rably reduced and the greater injection pr_sm_'_ and shielding of the fuel nozzle will help prev.s for li_t a:'rcr_ft.112 W_h the natural aspirating carburetor we have been discussSng. the system w_ll be quite ei'f_ctively protracted after about 30 seconds. that impact ic(_ and throttling ice az'.. aside from the fact that the pilot must remember to turn it on and off.. a_-ailable in sm_:. If fuel is injected _mder pressure at a oo:_nt downstream from the throttle._d._so are conditions that confront t_he inexperienced pilot. if the engine is throttled back and held at or near idling.l be apparent._n take-off d_sta_nce and reduce climb dangerously for short _nwa_y operation. This practice is too h_.nd. although the seriousness of their'-_ffe_._cnt exhaust heat._d for personal a_r_'raft.rsiz_.L and dangerous from other standpoints a me_s of ice protection. The common remedy for induction-system _cing. Th.'.d_. .fuel m_tnring _s greatly r(._ aSr flow.._not at all Drevcnted by thi:_ change in fuel _nJect__on. until rough or erratic engine operation resulted in backfiring or complete engine stopp:_e. available in all aircraft. ice may have interfered wlth the fuel feed.th_!s ropr_s.:nt stoppage of fuel flow. Coupling th_ hot-ajar control to tho thro_tL_ so h. But this is not _._.._nc._plete eng_ine stoppage occurs and power can no longer be recover.reas 150 ° F air is .l!_. better means of protection are n_od.i_d automatically _t low power._. If the valve is not frozen with impac_ Ic_ _nd can successfully shut off the cold r_n air..noug_.ts on air flew _.phazardous . Durine_ a long glide.-iscted when needed would help a groat deal. in some ca_es pilots have recommended leaning out the mixture excessively to force backfiring to blast ice formzttions loose._vel engines may !ength. One defect of this method of protection ..t. the prcsBure carburetor which has be_n st_uud:_rd cqu_ipm_nt on transport aircraft for many years end was rccontly m_. the induction system can become loaded with ice and later fail to develop full throttle power during takc-off because ice is throttling th. during the previous example. is that ftull hot air in sea. _. y t could be m_nua!ly s. the refrigeration _ffcct of fuel-evaporation in the carburetor body and on the throttle old._nually valeting to _u alternate source of air which is heated by an exhaust shroud. _ports _nd &.Lso b.!r._.:'s cannot take advantage of this improvement until mos.b!c_ ct_p tow<_n-ds making safe flying easier for th.]. Since the svpercharger "s heated.-l. A rad.._'wi]. The speed-density m..induction . are expected to _.tc pilot.ete sJmolificat_on of the ir_duction-syst m ic__ng _roblem lies in excluding w'-_t_rfrom th:_ on__._..on :[s now avrd. for the Dressuze carbuJ'ctor in th!'7 ]Z Some transport engines now :_mbod_ a f_rther im]:rovcmont _:n 9ressure-carbur.<.he _rcrt_cn on!. l_y the compr_.'_n%. The h.ly..'_bJ. rcs :arch at the NACA Cleveland .)__bie as decl_-_ssiflcd Advance R. se_. This is a ..incd l:_tcr a_d r.ng ms_1_s other than the csrburcto.:_'_dy been by in_stigatcd by N_A .md._Ls are prev. fuel evaporatio n is entirely eliminated..n_ . The como].urc for pilot _nforr_tion was prept_r. a Job the British -_re d..?h. sl_ow_.t_fy_ng to l.oir_g with internal hot-oil circ_lation.srn ti_at _:_.:w .ction system._d _ entirely from the vicinity cf the carb_ret<_r and fu. E_minating the venttu_i air me t_r reducers the smriousness of __mp_ct .:d i_rou..._rcr_'t._privr._ _%0 joint :fforts-of NACA and the Air Materiel Co_.arnborou._f130 ° F air is required form. The fr.u" makes of !_ght a. Air Forc<. Royal Aircraft Establishment.Laboratory in l_t_-_ 1943. Bureau of Standards and continued the.:. of natural aspir:-_ting carburetors.cha system an4 is currently be'_ng fitted to some' ra%hcr small cn_._d_d for properly proportioning the fuel among the sev ral intake p_pes._._n '_..at requirements for this system are still lower but ice on the throttl_s w-_!_ still require heating them. is st_.neclfied for +._. _-_ome of th . besan a broad investigation of induction syst:_m icing at the.llmcte.._sion of air.re a.._-tesone cause of throttling ice.rat.'l i_ admlttc_ through a _pirmer or thresh.st._teriug system dove!opec! by the.=:n_._. F'. Wat{_r _.'.co and e] imin.rch continues In along 1941 the these NACA lines.rovement in ice prevention c_n_ b_ accomplished by utl!iz..13 s. It is gr._. obtainc<i for the .-_d I_coerts while th:_ majority of thm results. Unfert_n_t._._tor fuel injection.s.e.:_dditiona 16 mi].].' for meterin@_ and then injecting the fuel i_to cr near th3 englne cy!indCrs._ this !n country and has already been adopted as standard equipment on re.r sound motion oict.i rome _.ndo.lbe expl_J.1 nozz_"' .. In :.ine air inta}:e in combination with the last m:untioned fuel-inJ.:.Long and comm. n._r.::sults at<' now _vai.r_ct.'"_nos.evel _ng_n.h_ the ]mpcll_r dircctly into th_ supercharg_'_ >._m moans _f -iudic_ous a_r scoop d_s4t-'_ h--_s:. A still further imy._imolified form of full fuel _njoct'.c 4ssued soon as Wart-ires B..'_:l reduction in heat requirements is th_Js made posJ_ble and the cng!ne is free from _cing above 40 ° F.-_inZ s_m_mar_zed in a Technical Note. xt.. The top left illustration shows the rel_.typical impact Ici_. Ninety-f iv(-.. water-excluding _ir intake. F_g_re i shows the ben.._ature is plotted vertically and water content of the air.ffoct on the icing l_mits with ].tivs paths of wat_.!o opening has little _. a complete absence of scr.ly reduces the temperat_u-e at which _ce . Th_s _s the first step toward a truly sheltered.njection. .asing power by opening the throttles ge.onfor a_ experim'.r clear a.114 system icing and de-iclng as seen through cpeclal transparent carb_u_ebor nl e_:_ine parts._re _ for thr__e power sc_tt_ngs. the throttles finally froze in the position shown and could not be returned to normal when a lsrge plus of ice brok._tally.gon the exterior cowling s'.and. For the purposes of this paper several flgur_s have been prepared to illustrate typical research results. water and air str_amlines aru she'.n ice tu%der _dentical test conditions._._ill fo_m in the system. Tlu'ctt. nera]._ntal form of undercowling scoop.rasobserved f. F_gure 2 illu_trates typical without t}'e carburetor screen.!r conditions at inlet temperatures of lO0 ° F. next the resulting impact Ice obtained _uring tests _n the Clevelm_d Icing Research Tunnel. It is hoped that thi_ f llm may soon a be dcclassiflod and m_de6enerally available 1'or pilot ii_formatlon. hor-lzo. incr_. with and pressure carburetor F_gure 3 shows serious throttl_ng and fu_l-evapor _tion ice in the sup_rchargor _nlet e. bottom.eyge amounts of water pr_sent. The limiting ic__ng conditions for this engine and pressure carbuy-otor are shown in fib. C_rbur_tor air tempe_.. ser_oas impact a throe-barrel Icing.fits of watc_r _9p_u-ation in preventing _mpact ice within the induction system of a large twln-eDgin_o cargo aircraft._ft of the_ central w_rtical llne ar_ clear-air conditions. n. percent of the freezing water was excludq:d by this design _rith no loss of ram in cruising flight conditions for critical altitude. It is _mportant to not_ that ice _. to the right are cloud and rain conditions (simulated rain injection_ in carbur_:tor air inlet in ga-_m_ p_r minute)..'r and air _nto the standard scoop entrance.u'fsoce. On the top r_ght. of with spi_er ._ loose and permitted %it flow in excess of cruise power... The t_ro pictures at the ri_ht illustrate the result of continually op_ning the throttles to m_inta_n power Luring icing conditions.:_.!bow (top photos) _md underside of a twoba_r:_l pressure carburetor (bottom photos) for three t_st conditions.. To the ]._. in the bobtom photograph :!co almost completely blocks the csrburetor screen e_bout 5 feet do_nstre_a of the intake. y with th:is sam. Despite the gen.n . The _2/pr:_.) '.. and Jud6_ne_t than co_. I should lik. It wou.imJts for the occur.oc.'._.ing char:_ct. upper temperat:zre l._r. silence a'zd backfire.-fr(_. ]C_:.r.o op.'..doption of the automatic choke . kll-woather us_g.r:_. _ _ pr_'z.t troubL.ssive _. It !s important to note that rotor of t. oD_. nt of _e_at_w.ot?ed vertically emd water-a_r ratio horizontally. that hijhly trained pilots_ adequate c_z.._ realized.]. alcoh¢.s the use of natural :_. op_r'-)_tln4.rbur_.:.-._ss. The six heavy curw_s represent the. _ven though first cost st:!ll dictat.s of const._s.b_r_tor h_at.clt'_mur._:%u..a of spSnner fuo3. h_xm:_dity (R. oper:_ted as saf_].binations.orai seriousness of the lc._ pilots._ful con_'_[deration. .l sprsys for _-_r_ency de-icing._d _y raill'q_is of automobiles.._._u private pilot. by p<_rsomloipossess'ng f:_r h-.Jct p_a'sona! aircraft to b_.tha5 th_.'.t cruise power for _ix different engine carblu'.u sloping u_ward from left to ri_ht repr_*e. i_nj_ction in the latter case.control!_& hotsDot. _'_'eaclose to the top of the hood. and freedom from bacld'ire dangers_ long ago caused the sutcmobJl._ right.ch prew_nt ocriou:_ icing during cold.:.. to review _ a prc'ced_nt ests.on system icing a.in_t_ comparable in siz:_ to thos.':ridhermostatically t . constant perconta_:'. The c_u'ved lln'.) from I0 to I00._orthy of our cu._%monst.h.ai_n p_.mprovemsnt from the topmost curv9 to the lowest curv:J _s accounted for by the us. starting _l_d at low _:ngin9 pow.st_ starting of with low volatil'ty fuels '-_n_freedom from the abuse or forg_.._ q_)i_tn.m does -p_._ ms_utm"actur.rat_.b]._ _id hotspot _ wh_.115 If you w__ll dlsre_Tard the co_lexity of f'gure ') you wJ]l note the t carb_n'etor air te_pel'atUre (C.!& not bu rsason_D]e _ to exp. Ths lines sloping downwerd frcm ]gft to right .tion.ish_d by another personal vehicl_ which has som_what similar problems. The dictates of freedom from gorasive wear._t by t.es cl_ar air con3_tionu i- -_ are to the left of the lO0-percent lisle. cloud #.sa_.E._ equipment bec::_use of th_ relative inexo_rienc_ of the av._ in mo_:t i_ght aiz'p]an.ora_..spirating ce._rlutics of these military and transport induct-'. and good :!nstrumontat_on insure safe.tors the bad ic'ng qualities have be.._d drivers led to the _. _e all r._tor co_.Ate the a_r intake in tho warm sheltered._re ]in.tom-_tic chok. screen combination._:ntday automobJ.T.aliz_. If you will p_rmit.tfu]_nuss of inoxper_onc.on systems it must h. This c_xs_%pl.rs to adopt th_ now familiar a!r-..A._Idrsln to th.tones of s_rJous _nduct:._nt enthalpy H._:ct_d of privat.zm_:.le induction syst.:_rat_on of cn_'.:. of automobiles mad_ it good practice to ].ss sk[]. is %.._neral a._d_:quately offs.only _ro._ oz_hibit serious icing tendencies in c]ear-a_r cond[tlons.h_ engin._. p_ will have been taken toward the _osl of f.y a_ a model to cop_.h_ ail-w_.]. Fuel A injection at el_ter_n<._ restudies: (I) (2) (3) A s_ng3e.md and the princ"plos ar_ understood.. a v_ry satisfactory induction system for personal aircraft will contain the followiP._hcr por._on<l aircraft. _holtered all" intake in a r:m_n_d p]enlml ch:nnLer.h_ carLu_ctor.Z .:_.device (4) (5) _ach bcdy in3_t _a!_e. warm.cing Sc_e b.16 I dc net hold up th_ automobile induction system ncccss"_r!l. The reward for such . It is possible thai:.vpe of speed t.% design will be . the cn_. ._Ino and simple warmed throttle by it.ir-clc.y. Intc_gral with (6) A the_.at low air t_mperat_me:._d Co extinguish dcn_It_ servcs to oi±cnce the backfire_. An ::.qostatical]y controlled hot:3pc_ to insecure good f_ei w_po_'_!z_=tio_._er unit which a_o '_ air !n_ke .._afer fiyt_ _d rcduc_d training time for m_m. future pilots and an important _. _d to G!{minate reed fcr FiAot attention to induction:_ys_m icir. repl'_. The elem_nts of such a _ystcm sr_ al_f _t h. H.A. R. G/M. 2695 in 15 minutes R-2800-51 power.--4 . low & W. 500 29°F.H.._'e 1..p-4 (D (i!) STANDARD UNDERCOWL Fi_. 31°F.. no reduction. C.T.- induction system at .© . air-flow 10095. air-flow Impact icing 10095.. R. NATIONAL COMMITTEE ADVISORY FOR AERONAUTICS Figure Z. in P. C.T. cruise simulated rain injection.' Effect of scoop design on carburetor screen icing. 5095 in 6 minutes . simulated rain injection reduction.A.. 40 60 80 100 PERCENT Limits 2? 5. 0 G/M 60 % cruise power Figure 3. 100% Water. ADVISORY FOR AERONAUTICS NATIONAL 100 8( CARS AIR 6O TEMP.00 ? lo. R.T... 350 G/M C. 100% Water...© o9 C oA..A oT.A..oo 7 SIMULATED G/MIN inlet elbow RAIN INJECTION of visible icing in Allison V-1710-89 and carburetor with AN-F-22 fuel..H.. Water.H._LOW CRUISE COMMITTEE 20. 37°F 100% 650 G/M C. 40°F R. . .T.. 0 RELATIVE Figure 4.J . 40 / _.20 _.Typical icing in V-1710-89 induction system.H. 40°F R. 006 .o °r--I (1) 20 70 60 CA. LB/LB Figure 5.020 TOTAL WATER CONTENT. oF 5O \\ 35 H.010 .001 . BTU/LB _30 4O 32 .060 DRY AIR . NATIONAL CONNITTEE ADVISORY FOg AERONAUTICS .040.'[.002 .004. the conclusions of this study are net directed toward recommendations for the adoption of the new engines. The new turbojet 8r_inc is less complicited thmu the large reciprecatln_'. safety. The _%_s turbine driving a propeller hc. The performance ch__ractorlstlcs considered are fuel consumption.sa fu_. e_d noise. an application of the rs_n Jet.hi<_" high speed flight conditions this new engine has a lower ratio of weight to thrust. i '/ The study gives an analysis of engine perform_ice but neglects other important considerations such as cost._en made to provide a b_._os_ very high fuel conmLmption._re _was estimated from component efflciencics observed in toasts of the small turbojet engine considered in this study. This fi{. Improvement of the . Th(. Figure i shows a comp_rison of the specific fuel consumptions..7 pounds per horE_(. This virtue Is likc._tenod in the engine as a power past few years by plant newer types of propulsive systems. explosion ram Jet has a ridiculous fuel consumption of ever 12 pounds per horsepower hour as compared with C . The t_boJet engine a]_o h. Those engines arc considered applied to _'n airplane weighing 1200 pounds. Sanders for 'i _ • ¸ The dominance of the _ ciprocating aircraft has been thre. The reciprocating engine is shown to have a bower fuel conscription th.1 concumption of 1. <_ngine_. Consequently.ly to hold this engine in a position of great importance for a long time. _nd :.w of the characteristics of some of th_s. No consideration is given to the possibility of improving the reciprocating engine.._powor hour for the reciprocating engine.. aircraft range. which is far sidereal in this prelSminary from exhaustive. the buzz-bomb engine. engines conthe turbojet ongil]o. and h.'s must have the seine static thrust equal to that of a selected reciprocating engine and propeller.:. includes of the.i17 PROSPECTS FOR NEW TYPES OF PROPUI_IVE AIRCRAFT SYSTENB By FOIl PERSONAL John C.__nany of the other engines.vingcruising an_ maximum speeds of 95 and 105 miles per hour with a reciprocating engine. and a gas turbine driving a propeller. and maximum speed. thrust-wei_ht ratio. The list. study.9 new engines has b. 3 pounds per horsepower hour. Since the qualities of light weight __ud simplicity are sotkght for personal airplanos: a preliminary rovi(._sls for a scientific guess _s to whoth_ " research on th_ application of them to personal aircraft should be conducted. A b_sis for comparison is achieved by specifying that the several engin_. _ccounting for bu_nqer and windage losSes. which corresponds to c. Estim_t_s of the porform_u%ce of such .7. _lo_ing for the _ui_t of (. An engine ¢. eou_lly The reclprocatin_ as he_vy s.t order of _:_erit th_n found in the study of fuel consumptions._d_red._ per ho_.8 components per may ¢ventuK_lly hour.. . In this engine air is induced into the hub _nd p_uped by centrifug_l action through hollow bladce where heat is _ppliod by combustion.ire 4 shows the effect of o.1._8 efficiencies ._.. In tb. !_ll_. This value is compared with the ftlel consumptien_ of the other engines in fii.._nj_nes cons. postp¢nod until must be considered.re sheba] in figures 2 and t_Ja'boJet ong_no cm(l g?_s install_d in a porson'_-i 3. on the fee" consumption.'_teng._ircrsft _'hich has been suggested fr_quently is to attach r_ Jc. At :_ M. _nd the hot .7 co_ld be realized at the r...fuel con:. exprezsQ_d as Mach number. _ud lih_t those are the The turbo._ch number of 0._Jot._ir is ejected from the bl:_do tips.13. reduce the fuel pounds horsepower A_'tists' impressions of how the turbine driving a propeller would be airpl_z%e 8.s the engine explosion -_-ith its ram Jet propeller is (or buzz-bomb engine).ure 1 arxl is shc_a% t(_ be much higher than for the reciprocating _ngine.__ne was obtained from the actual weighing of a turbojet engine of the size used in this e_nalysis.no is somewhat two heaviest llc_hter m_d .._. Much n_ber A comparison of the ratio of weight to static thrust shows a differn_:.t speed of lO0 miles per hour the theoretlc_. indicate that _ fuel con_t_%otion of 3 pounds per horsepower hour might be expected. Then Mash numbers _±bove 0.ts to the tips of the propeller bl.mnptlons will be achieved only '_t sp_e_ of 0. Fl&.. s.ngine_ are shown be be oqua].. The we.. Th_ buzzbomb weight was obtained from the thrust-'_oig_t ratio of a buz"bomb motor._procating _ng!ne The J_eig_ht of the t_rbojet eng.n_bl_. now Discussion because the of the rcJn-Jet engine h_s been char_. A plan fo__ utilizing the ram Jet on slow-spool _. th_.nn origins.irspced..'_not with the airplnne J traveling at only lO0 mile.l fuel conmL_ptlon Is shown to be at the very hi_ flgtrre of 80 potunds per horsepower hour.ctel_istlcs of this engine in some detail. i_ comparison shorn in figure 6 tl_e static thrusts of the several c.:aringend Dropoller.s.mbodying a similar ides is she:he in figure 5.consumpti(:n of these to C.zd_-._bove ". Reacc._'_ht th_ _as turbine of and propeller' was estimated from the weight of the turboJ'ct engine._.opcrato_ oropeller is much li_hter _han the rec. t_tor to the roc_prcc:.tno ei_ificant differanc_s should be pointed out from this fi_'.lifications as to the accur_cy cf these c:-_iculations uust be m_. which produce the s:. The weig_t-r_ug char_tctoristics of the o_p!osion. -_u 7a.t_d msx_:. When.i_uq_. engine _ire shown in th_ top curve.. propeller..r is here _.r enli_inos an& %he fuel ].::nv._ because th..:ight-rcmgc b'_si.) s!z'pl:u_-o_. It m_..7 high.. TM.._irc_2t. better than th_ explosion rein-Jet engine but is heartier them the reciprcc___t ing engine.ncsequipped with the Jet englnos hure:[n described would be greater than for the propeller-d±-Iv.Iz_s.. Ill this =" _" • ig. .h fuel conmwlptions at fllg_t spee&s below lO0 miles per hour.:mesb._xim_vafly_ng speeds of the :_irD_._t. The g.[propeller end igs fuel would be no gr.:tr_me!._mon ter__.o. Iittle c_n be said to recommend this t_pe of _nglno on a w.. the -.ybe seen _Da_ at a r_u_ge of 150 mi] t s the' weight of the Jct-o_ereto_ _. _-ut_rc pow.-t_.___iz_].¢light f s_ee_s using thusc . Inclu&Ing ez_Ine._ fuel reqtJired for specified flight di_t::. considerably _xcJed the woigho of thv remaind_)r of the .._.ct.nd the sotu_ce of th_ difficulties b_vc arisen from the very low static -thrust and hig.qnglncs.. f_r rani_¢s of 300 miler or greater._to be a close comp_. Figure 7 shows the resttlts of _uch computations.i19 The two characteristics Just considered of fuel cnnm_ption /- - and engine weight m. .. _i_uro 8 shows the cutim. .ule.. :_cce."_tic thrust _:s th_ propeller engines. Thus the . The low wci@ht of the jet'o?er_tod propeller without fuo! opens a y_essib.. Jet-p_opulsiou engines _y_ _(_d.:.c_ _quired is e:.ti_%! weight is approxim.ei_t of the.r_._:ed to i05 miles per ho1_.n.tt the exhales!on-.q.ting encine._de as before.n(__ _ the turboJti. It m:xy be seen tb.<_-turbime <_rivcn pr<:po]h.r for the three propeller enQ.%y o_ reduce& to a.g_:_In sho'._c._._tcly eqt_ivalcnt to that of tb:_ cbhe.Jet ])rooulsion cn._. The turbojet c_inc i_ _?.._e the ahs:z!ssa is the r_<_e for _hich th._ the Jot prcp_ller weu?._n :.ncos. fu31 load of fuel on bo:_Jd.__mputlng the weight of the engine plus th. The s_zae qu'._-Jet . :_nd fuel to_uks with fuel. In _ho cache _of the reciprocating engine it may be seen that if the airplan-o were dcs.-'r than theft of the rcciproc%ting oncinc a_J.dto fly 300 miles and had c.zinu would achlove a f!l_. in f_. the wo_!._sorie. however. achiev_ !42 miles per hour as conpe. The compavisons so far mqde have been quite unfavorcfolc to Jet propvision s. The p_rform_nco of these two power pl_ts is eutims. for short:_r flijht._ in_..bly greater.:_ ..own t< S_c._1_tof the sower plant would._hb. soecd_ of 160 miles per hour .d be lighter.by c. cc.s desi_ed to fly ea_d the erdiru_te is the weight <_t bake-off of the entire pom_:r s_stem. the _ower _ail_hle _t cruisi_4q _ec_ w_th the Jet engines is cons_ter_.ility for its u:]e in spi%e of its hi__ fucl const_pticn. tod to be suff_cicnt?y near equaliby tD._ • system would be 240 po_nds.. ::ll_r _-_rrant further stu(ly and research.._. In clo_ing.f encoura_in_ u_. im_ractlcality cf applying them to the _:_rsonal eirr_me._s_arch .t.. it ap_ears [...rag._'ill not be ._ coeffici(_.hat whgre range is a factor the reciprocati_:% e_l_dn._.t D/qS d... !is chief compstitor app_ars to be the .. The gas-turbine engine appears to be a close comFet_tor in weight and fuel-consumption characteristics but I hev:_ no _dea what the comparative cost wculd be._s that tile _au-turb_ue drlw_: propeller and the J_t prop.120 In cast.]c engines _u ]_t ai_'craft. sc_ engine_ _. fi6mres arc estima_es r_ade for the purpome of _et:n_ninlng whether research on a'_.ing up the accounts for the seve_a! engines discussed. ho_¢ever.od propciier.s-tu_bine driven propeller and the Jet-epcr_. I would like to caution you that the_._ :_.._ should be cons_dered for th.. squ_ro pounds feet ._. per square foot . Such a recomm<ndation._:i'ul for long-range flighgs but its sl. to research organizations.a! the .mpliclty will tend to r_ake it a cheap(r eng. since the _dertak_ng of the dcvelopment of these engines at present would be extremc]_ costly a. • of the._ne_ thus meeting one of th. Th_ _et-opcz_ted oropeiler . wing area. SY_C_O_ CD D q S •dra. is made on. pound d_-namic pressure.may rew.._s..q_" further r_. Jet proio_ls_on might b.. The study inAicat._ usclul for high-speed airplanas with short ranges.. purpose ¢.: requirements of the future engines.'!l be as good as alkv of the others._ of . Airspeed. fuel consumptions 100 miles per of several hour. power L NATIONAL COMNITTEE ADVISORY FOR AERONAUTICS Figure 2.- Turbojet-propelled airplane. .- Representative systems.SPECIFIC O 2 I FUEL 4 I CONSUMPTION 6 ! _ LB/THRUST 8 i HP-HR io | 12 EXPLOSION RAM JET TURBO-JET JEt" /'_ • OPERATED PROPELLER GAS TURBINE ¢ PROPELL'ER 1 RECI PROCATING ENGI N E Figure 1. NATIONAL CONNITTEE ADVISORY FOR AERONAUTICS TEMPERATURE IN BURNER 3OOO .- Gas turbine and propeller arrangement.3 .- Steady flow fuel consumption.4 .Figure 3. .9 NUMBER ram-jet LO Figure 4.7 .00 (OOR) RISE i I I I I I I I I .5 MACH .6 .8 . G ! . THRUST DRY o .- let-operated RATIO propeller POWER =' installation.4 ! ./ NATIONAL COMMITTEE ADVISORY FOR AERONAUTICS Figure 5.8 i EXPLOSION RAM JET "l"UR BO . SYSTEM STATIC WEIGHT.J ET / JET-OPERATED PROPELLER m GAS TURBINE ¢ PROPELLER RECIPROCATING ENGINE Figure 6.'Z I .- Ratio of power system weight to static thrust. . pounds. PROPELLER 1105 RECIPROCATING ENGINE 105 Figure 8. (3.. EXPLOSION RAM JET of several drag. NATIONAL CO_HITTEE FOI_AERONAUTICS .59 square feet. 372 pounds. Equivalent value of CDS.(D [.JET 14"?_ JET OPERATED PROPELLER n 105 t' GAS "TURBINE _. ADVISORY 37Y.Range-weight comparison Static thrust. MILES 500 I I I I J oCNJING u.Q / .I U _z Figure 7. 153 power pounds. G_ . systems. 0 _D >. 95 miles per hour. -Y _. 6._ __-_ OO "r" C_ _ 2o0 _1 o 0 I00 200 500 400 RANGE. cruising speed.Maximum speeds of an airplane with several types of power systems when all power systems produce the same static thrust. static thrust.o ! • 13. cruising IGO MILES PER HOUR TURBO ... u_ 400 u.v SO0 n. 12! < NEW RESEAI_C_ -b _L . _ of the large number of variables that can be changed in designing the airplane for good stability charactcristic_4both l_ower off and po_::er on._sthu scope of the projects has been extended to make +. .NA_C. A similar study of latersl st:_bi!ity is also being made.arci_ on a wide variety of oersonal-aircraft problems.from the standpoint of longitudinal stability and a!_o to det._ results of the rcs.later._l industry anti Go_ern:m=nt organizations a large nu_oer of suggestions for rc::. bei_. These data vdil be analyzed to dete_Jne the better co_?iguratio_.. During this cc1_ference it iia__ been shoran that a oJnsiderable amount of the Committee's research on military-aircraft oroblem_ personal aircr_f-g. and in some cas. by anal.. I wd.ac which of these projects arc api'licable t_ personal-aircraft problems.ll discuss very bricf!y. it is extremely difficult to aft'ire. therefore. in this case.h. r_sults of which are now becoming generally available_ will a:[_]in providi_g solutions to the problems submitted. _..v%ical means. One of the ._.)_V.. at. turn aimed direct!:y at the personal air_?_ane. in addition. is of direct value i. vS_th a larg_ part of the work..study prob!er_s which are peculiar to the light airplane field.. the cmvrent research in various of interest in tkls connection.. of a!l. INTRODUCTION The NACA has received from sever. The Langl_y L.ze research suggestions we have been interested. field_ to the which is _rojects P_:SEARCH I[_VING APPLICATION TO PERSONAL-AIRCI]AFT DESIGN Stability % ._ the significant factors that contribute to good stability .. we have examined our current and cont¢.. the move favorsble configurations for stability. v_l].<.mp]ated research in wzrious fields to determi.md.g done by t}_ Amos L_bo-.h_racteri:_tic_.n the d_s_gn of Next._boratory is. making a collection of the data of the many powered . in determining the extent to which NACA wartime research. several investigations h:_vc been initiated by the Committee's laboratories specifically to . OF l_m_Dv_m P_R_.: In vie_".-a[.nd-tunnel models tested during the war. first.s .ZO2 CL_RE_Yf AND PROPOSED IN NACA PESEARCH I_gUESTIGATIONS DESIGN Jr'. R_RA_T By Harrison Chandler. firs +. In our study of t]:.2arch mor_ direct!:l app!icao!e._rmin. .tenl.....tallied to :'t.....ion to the changes in stability that.t!_-c..L tunnel air:....:)_v-turbul._tions of :<pc.... T ..:s.)lem is '71:it regarded as particularly important for standpoints of both safety and utility...... ng d_h<...ane model. 'F]Jis [:. Oar curr. u_ed in c. high-lift is an im)_crtant problem in the design of light airplanes...f a flight study with a represenLeti"e frc_. < Controls _.ithparticular ai...] ...t.]t...-riigb........ _'. ..]ral... occur v..r< sc.n bh_ ]......... _Ying Design The Langley low-turbulence tunnel is ext.o].. At the AmesLabora+.._search is directed towsrd the oroblem of ro]!ing momentdue [. and (h) the effect.......hen pc.. In order on d..123 portions of the AmesLaboratory's lateralst?td!ity r.. In addition to this research..:siip.of high-lift flaps on lon@itudin_i and late...l data for airfo..ils fitted with a ..n6ine tint ".. . These projects are beinc reported individual3¥.....L_acterlsties of airfoils to a low r._vic....... as Mr. As you knov. appear to be the best solution for this prob]........'..... v. A_. igat.inveU:i.wmmic condiIcting to assist in reachin_ an unde.id_ variety of types of high-lift c._.ynold. taking into account the effects of flaps.:-'rc_......._njunction "_i±...............'..... problem of obtaining hil/h lift.....ither alone or in combination with the so-called ".. .... (2) the effects .... ..s the problem of provi(!inc suitsble lateral-control devices to be... i._T-laycr control as s m_a.!ide-p::....-al stability.dirlg th_ mr_ximum lift of airfoils..has made some specific tests de:dAng vri...)f.er is applied..gn data which v.iL devices and flaps hav.._n_Z its iuve:_t..r_t:n_din_ ')f po_ver effects is longitudinal stability the L_ugle......ub stud_T deals vdth the effect of length and spanwise loc...:id...v h_s b<......ing minim1_-_ effect on airpl_. d and other contributing factors..q. Spoilers.Lge....eu initiated to provide a design basis for r......._ of numbers in order to provide des_.orya p_'<>je.......th (I) the ef?ects o? tail length......feeler" ai!e1-'ons..he effects of wing po_ition..o ..ne tr%m..._!_ncc tunnel on boundar._::nd..ntr..g_tion also is under _'.ayin this tunnel to provide ...........h the high-lift devices..n_ for cx...:'ng....on of th_ clm...rc'. personal eircraft from the As a pert of its research program on stabi!it:/ problems the Langley Laboratol_y ..... ... T. work is being cerried out [.... i]ers._:u(_ -]:_.! m.'...._ll be directly applicable to p_rsonai aircraft over their entire speed ra.. ..._ngine per:or...w (3) +. I{....n_tioned..em..l!ied _ith the. . that the engine performance.c__mu.7.. The Co:mait+.. invcst'_gatior.. es The Langley: Laboratory is cont".zstems._.. forms of propu!sives_.stigati:gns in ?la n t s t..or ._ . . progress _n the craft (_esi_n.qtcre:.:. .....ngin_ using i v._f new _..._'.cv.A...ct.as .n One ef these _rojecis.u'ther pewer-plard._c!ude..:of_..-_aii hu]l.!_r advance r......i!i p.. projects of the:_e _moject'-' . a stud.."...::.y ":%_._.t i..-_a-'ch i._' of i...an_ l'-nl]s.-_:_ch d_:. ".cab]e to pro:...norder to be fully a:_?lJ. the results of this work si:o'_:cd.erodyuamic tests of hull models i_ a 7....:. • NACA of which were prepared in this research have been found to re...._impe!l_r in the nerma! high-test supercharger.roble:'as peculiar to pcrson.&oor_.Jcrib_c]by "....ir- .!D-dro_Yum_._orkv.c _&CA _ _ "" _"_ . gasoline.de@ th'_t research . hov..--_scrib.her studi=s of the _:la_:t:_.c characteristics body forms of pl._uir_.'..er work in the NAGA bowing tanks to st ud.. Sanders and: bhecefor_.resently discuss intercept to personal aircraft.r..terest in st'.ic research studies on the performance of seap]... hyd_'odyn_zH.'.l.f propc]." .. and a sy_:ter_:a'5i.. Sev_ra! _mch projects ar.ft.. This . requires no further claboratio:) on my part. Ir._d _rl a paper /. .:n revived since _ ' be_n recommer.._.ic characteristics. . .l.currently i. Seaplar.s r_._s as fur!._. of the order of 0.as well n s furt_. extension to a lower _a_ge ..*. Crigler h:_s outlined in his paper the results of rcceut research _lich are r_g:_rd_d.h:}: field of pov_t_r U!arts nov: it.._ .n progrc'_ aud inci_Mc :-x_ch ". oersonal aircr::...by lO-foct u_Lnd tumn. engine-perfo_auce te:_ts _ith a safety fuel. .. w_t'._. [.a_ing surfacc_ hulls. r.-_. has just bc._.".eccntly z_ade._." _he p:_opcller selection ch:_rts propellers of high _ ff_._-..ral ..:irf:.ee':: Clevclai.fet:/ fuels has be:.cbion _ -_.te:.._. v_'. '_.. an inj_ckion-typb . a!:en.io (V/nD). This ._ this l'roblem be carrix:d _nd it out to has / provide a better basis for 'ace of [:afctF fuels in personal-aircraft engines.ud7 of -_.. the _. Parkinson in his paper.asqual e to that of the _.ill be undert.c_._Orl_.:1aircraft.ftc_rs of flv:'ng-boat Fower Several inv_.'.s directed towed t_.. its i:. as mo_t suitable: for the scl.::ver..d Laborator.l..Propellers _r.f ha:. _.]. :..[ts of !-. and engine-induction system._Ircraft.'_' means of icing protectio_ for all _irplane cor.n:_tr_ments.c_nt s_._a:so_L:A.cpemda}. aircraft arc for all practical purooses not designed for . No!house ha:_ illu:_trated in his paper a orit_:rion of relatiwq. i_izcd in a.. for r. was based primarily on re_u._nd amphibians..h_.@5_ 'ati°ns .Idition. study of meea%s for r_kil_ personal a_rcraft of 3pirm_ng evt.Icing Although present type.'. c:_rrying t out a nm_ber of investigations or. Rc_ar(_h is being directed crit. a Sp trming The first charact_ristics of these proJccts is an of persona]. an(_.'_e typical light a irplene conf}.m[:_l_tc_. From the results of the _:ork a!re_dy cr. thc._mple application for _. a]so._pln recow_ry° Thi:_ criterion.and foul. ircraft of the future a must b<:capable of _oth fair.'./ to consider that per.s currently directed toward to touch briefly noon the several r.Jperat_on in incle:_c_ff.or_"ig_r_tions. .1. %.it csn be stated in ge:.f .lponent:_ inclu6h_g the wing and tail surfaces.sbeen fom_d to orovid_ the most effective and :i._._rion for _. of course._ of persona].e:_e_:r<. wcat_''.erai that the use of ha-at in vari. propeller. i_ nevertheless necessar. twin- are oxT.afety.ind_hic]d. being In _.itary alrplanc:_.y ].ncs._<_rsona3_ircraft._.. I now wish tior'.:_ona!.n if stalled. radio ante_as.h "r_vestigain orogress J_n th_' Co'_itte'_. a. it.ow weight as comlcared to the wei_fhbs of other anti-icing and dc-i. tbe protection of aircraft against icing. it is bei!cved t._the_"_a! ice-proLecbion t eaui. which tractor airp].irg systems.weather flight in order to realize a satisfactory de[ree of util'ty. izr_cstigatlon of spinning _.'s Lab(n_at°ries which are solution of problems peculiar to [..r.om_mt for aircraft can be desi_Tned. incapable ..tth_ pr¢._! boom pusher airplanes.s he mention_d.m_ f'or_._ctod to i_clud_ low-._s_arch conducted on mi]. i.vor'_:ltypical tim(. The Amesand Cleveland Laboratories at the pre:. tow_:rd verifLcatlon of this light airplane _. ea_d high-wing ta._re ut.y s.:_ Ls.._ntime are. i "s .inz for a large _ircr<:ft _njJ. is contemplated tl%at this rcs.dn this inv.d _..:_'rest.stii _c.-d_'ag apd com'entional t!pes of airfoils are being i.-em_:iag sd.or - %_._ cooling sy...cn._ in%.oind._s_ including r_duc%. vices under actual o rJcrating zonditions.d!rectl._. J:J_e..leer Laborzt.icity o_ %he _use_ce 0£ moving parts_ low _velgb_.e_tJ_aLions of %he ch_ract:._u. Jet-Ejector The Cievel_nd Cooling of 2ersonal Aircraft Y.rn<s'_ d to in!uce flow of cooling air thro<_gh the cowliug around i.agated by the propeil'_r a_.rp_et.'<tion.!y has condu(_ted r_..it{'. _ the anti::'e :rip.s investigation.i.h on Jet-ejector coo]. and i9..k::b-oj_.. c:_%e 'neous _or reducing 7_rope!ter moise.aust and Drop_].uginu..ora_ory recen+. The _.t['l vio_v tov.'rJ. Both i_h. f wj!_d-tunn.s applicsb'ie to %] c desfiga of pcrsoral aircraft v. Airfoil s As indicsted orevlously._... pa.Vogoley._pers presented by Mr.dl!be C/..-_t.x[ out 1_3"the s gudy the applbcation of 5P.sigr_3icpl.utions to 5Pe problem.souud ]._e to mea:_ur<: the... te:itio_l in this research.._.})r.lu.':_.anncl is extend[rig _%._ypresent in bh_ exhaust _a:: is h:.rd a pursuing .:.-'! personal a irp!a'.s of t]mt research in tt-J..dn:_s eve..ris_.. d._is c.q:._is of the v_rious do.(!the engine _xh_:ast of airp!anos and_ secondly. As the Langley L_. inasmuch as the propeller has _.de of the noise prop. . ory to is an inves%.5J.] be conbjmpd.1%esbs o on a_r_(.:s over ct]_er eng_nesys/e_.:ud to he the v._rricd out at sm%il scale.ductJon. in this m._m !!o...:ill h _. '!.-o"st offender.":e re.xp<:ct_. Regier and i_9".'.l data which .tic.rc:]_ Co determine the magr. is planned to utilize the resu).". This research wil. I:.ig_tion being (:a.w.126 }_oise Reduction Th_ s_cond of these projects is bhe resea::ch on noi._ector ..-:sults of this resea:'ch arc also _._or.ion in coolin%' drag. hc e. !:oth :51_gi._larly _7J.seal range of inb.s. indicated in th_ p'.t' I. .>'.ount of resea_'_.further the mcr.arch wil[ include tests on a %_Tic._]. The NACA Latorabories have carri<'d out a certain m_...ob< of o • _ "_ value ih the i_t_.m_i_c:{ fourth pro.LOl_o to a:i"_ "_ '" a low range of Reynolds numbers in order to pr._.. _e_in-7 because cooling appears to offer several advartag.ne insball.cl.ne exb.).'._: .ve:_al.lcr noise "_. the LmIc]ey low-tur!_ul_:nce t.n_d possibly some noise iampening.i.ido _Jrfcq._%hoc of engine ceolJn-_ the <n_. lo'..u£c:n on per$on_l r ircraft. '. which is nowbeing oz..Icing of Light Aircraft Engine-Induction Systems The last of theso.thgr_.ganizcdat b_. and one that is regarded _.projacts. Hm_te'-" as already indicated in his paper the rature and ob. Wil!_on i. p_ • i .iectives h of th_s in_=sti_at_on. is an i_vestigation of me_. Mr.ns for p_otecbion of !i{_ht aircraft engine-induction syr.tems against icing._.at importance.eClove!and Laboratory. Documents Similar To NACA - Industry Conference on Personal Aircraft ResearchSkip carouselcarousel previouscarousel next51440 DB-110 ReconnaissanceFeasts 2012 Are U Ready? 2011-06-17-DP-295-AEW-DATA-POSTER-GEN-ENJ.M. CardonaNASA SMRC TechnologiesGod's Holy Days vs. Man's Holidays_Which Do We Keep?City Under the SeaTischendorf. Novum Testamentum Vaticanum. 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