6th Responsive Space Conference AIAA-RS6-2008-5003Investigation of Combined Airbreathing/Rocket Propulsion for Air Launch of Micro-Satellites from a Combat Aircraft Avichai Socher and Alon Gany Faculty of Aerospace Engineering Technion - Israel Institute of Technology Haifa 32000, Israel 6th Responsive Space Conference April 28–May 1, 2008 Los Angeles, CA A ducted rocket was chosen over other ramjet configurations for its higher thrust coefficient. Israel. KEYWORDS: Air Launch.il Alon Gany Technion – Israel Institute of Technology Faculty of Aerospace Engineering. most designs consider heavy aircraft and launch vehicle to place a mini to a large satellite.AIAA-RS6-2008-5003 Investigation of Combined Air-breathing/Rocket Propulsion for Air Launch of MicroSatellites from a Combat Aircraft Avichai Socher Technion – Israel Institute of Technology Faculty of Aerospace Engineering. The concept of air launching of a satellite from a carrier aircraft is not new. coasting time. Optimization on initial flight path angle.il ABSTRACT This work presents the analytical results of a parametric investigation of a launch concept of micro-satellites from a combat aircraft. Microsatellite.6) provided to the launcher by the carrier aircraft. Haifa 32000.net. assembled from a ducted rocket (ramrocket) 1st stage and two solid propellant rocket stages. Ducted Rocket. Haifa 32000. and the launcher will be a 3-stage vehicle. and ducted rocket sizing was done. It is demonstrated that an air-launch of a micro-satellite from a combat aircraft is a viable solution. however. for various missions. Ramjet NOMENCLATURE A A2 A2a A4 A∞ AC AF B CD COTS CT4 D DR Cross section of a satellite [m^2] Intake area [m^2] Air entrance to combustion chamber area [m^2] Post combustion chamber area [m^2] Free stream intake area [m^2] Gas generator burn area [m^2] Gas generator exit area [m^2] Ballistic coefficient [kg/m^2] Drag coefficient Commercial Off The Shelf Thrust coefficient at engine station 4 Drag [N] Ducted Rocket g GTLV h LEO m MSLV ORS P r RE T USAF V Vc X F Of the fuel Gravity [m/s^2] Gravity Turn Launch Vehicle Altitude [km] Low Earth Orbit Mass [kg] Micro Satellite Launch Vehicle Operational Responsive Space Pressure Radius from earth's center Earth's radius [km] Thrust [N] United States Air Force Velocity [m/s] Circular orbit velocity Down range distance [km] 1 AIAA/6th Responsive Space Conference 2008 .technion. The carrier aircraft will be an F-15 fighter.ac. The solution presents a concept for placing a 50-100 kg micro-satellite in either a circular 250 km low earth orbit (LEO) or a more elliptic LEO. An air-breathing engine provides much higher energetic performance compared to a standard solid rocket motor (higher Isp and lower mass). 972-3-5325283 asocher@netvision. Israel. typically launched today via groundbased rocket launchers. It is the authors’ intention to present a method for air launching of a low-cost tactical micro-satellite. Documented air launcher designs usually incorporate a lift aided trajectory. on demand. using a weight economical vehicle via a Gravity Turn Trajectory. The option of an air-breathing engine for the first stage results from the high initial speed (as high as Mach 1. 972-4-8292554 gany@tx. (3) and (4) into Eq. τ = ⎢2π ⎢ ⎣ ⎡ r3 ⎤ ⎥ ϖ ⎦ ⎥ Incorporating Eqs. The amount of fuel required to gain the said ΔV is defined by: ⎛ − ΔVi ⎞ ⎛ ⎜ ⎟⎞ ⎜ g ⋅I ⎟ ⎜ Δm f = mi ⋅ ⎜1 − e ⎝ SP ⎠ ⎟ ⎟ ⎝ ⎠ (6) INTRODUCTION Investigating the sensitivity of the lifespan to the satellite's cross section and initial orbit altitude. The satellite's lifespan is dependent on the ballistic coefficient of the satellite B: B= m ⎡ kg ⎤ 2⎥ CD ⋅ A ⎢ ⎣m ⎦ It can be seen in figure 1 that at a 290 km circular orbit. Illustration of the launcher's preliminary configuration is presented in figure 2. For a microsatellite with a 0. one learns that for low orbit altitudes. Due to those payload characteristics. there is high sensitivity to dimensions and mass. and the efficiency of a closer-to-horizontal launch. The change in orbit's radius is defined as follows: Δr ⋅τ =r revolution ⎡ρ ⎤ =⎢ r μ ⋅r⎥ ⎣B ⎦ (2) (3) Figure 1. 290 and 250 km. a once-a-month altitude motor boost is required to enable a 306-day operation. A small enough microsatellite (in terms of mass and cross section) can be launched into a desirable low earth orbit (LEO) that can support effective lifespan for various missions. mass of 75 kg and fuel amount of 10 kg we receive the following results shown in Figure 1 for two initial altitudes. Orbit Decay vs.25 m^2 cross section (like a 0. it can be launched via air launch from a combat aircraft.AIAA-RS6-2008-5003 Z Δr ρ τ m γ µ Momentum function Change in orbit radius Density [kg/m^3] Fuel consumption rate [kg/sec] Period [sec] Flight path angle [deg] GME [m^3/sec^2] from total energy calculations via a Hohmann maneuver. The red diamond in figure 1 marks the fuel end point. thus creating a tactical responsive capability.2. especially if a high acceleration is associated with the selected propulsion technology.5m cube). (1) A typical value for CD is 2. (2) results in: ρ Δr = 2πr 2 revolution B (5) The additional speed required to return the satellite from its current position to its original orbit is obtained 2 AIAA/6th Responsive Space Conference 2008 . Whitehead [8] also showed the benefits of air launching via a small launcher into LEO. Time (4) CONCEPT AND PRELIMINARY DESIGN The concept presented in this work is a 3-stage launcher consisting of a 1st stage of ducted rocket (DR) motor (also called ramrocket) and two solid rocket COTS motors (STAR48V and STAR27). concluding that launching from high altitudes can significantly reduce the practical size of launch vehicles. This work complies with his findings and recommendations for air breathing small launchers. the F-5F Tiger II with one-half-size Pegasus XL. The payloads were 36. Launcher Schematic Configuration The launcher is carried under the belly of an F15 fighter aircraft as shown in figure 3. a level flight or a moderate ascent is required. By that. the launcher carries all the propellant and oxidizer on board.2m long Gravity Turn Launch Vehicle (GTLV) launcher that could insert a 75 kg payload into a 250 km orbit from similar initial conditions [3]. Figure 2. [2].AIAA-RS6-2008-5003 Boltz [6] investigated the use of scaled down Pegasus XL for air launch of microsatellites from various military aircraft like the T-38A Talon with one-thirdsize Pegasus XL. Figure 4. Launching at a high initial flight path angle as proposed in [1] & [3] is not applicable since the launcher will pass through the atmosphere too fast and will enter a too low air density level for an air-breathing engine before accelerating enough. Pre-Launch Launcher's Mounting In order to simplify the solution.7m long launcher that could insert a 93 kg payload into a circular 225 km orbit [1]. What unifies all those concepts is the fact that they all use solid rocket motors. 122 and 289 lb. respectively. and the F-4E Phantom II with two-thirdssize Pegasus XL. Estimated Specific impulse (Isp) as a function of flight Mach number for selected engines employing hydrocarbon fuel (figure 4) shows the advantage of a ramjet over a conventional rocket [7]. Among them is the F15 MSLV with a 4500 kg. 6. 3 AIAA/6th Responsive Space Conference 2008 . Mach Number for different Engines Figure 3. the use of a Gravity Turn Trajectory is proposed for the post-DR launch sequence. Savu [4] analyzed the launching of an 800 kg rocket with a 10 kg nano-satellite as payload. Since the first stage of the launch passes through the atmosphere. Past investigations of air launching of a microsatellite from a combat aircraft focused on all rocket. from a MiG-21 military aircraft into a 116 km orbit. we can use the oxygen in the atmosphere for the 1st stage via an air-breathing motor. A similar launcher that operates entirely via a Gravity Turn Trajectory is the 3900 kg. Isp vs. 6. Therefore. when using an airbreathing 1st stage. 3-stage launchers. It is assumed that the configuration can be installed under the F15 belly. The following figures 5-7 are for a constant altitude of 11. a moderate 16° flight path angle was chosen for the DR motor stage. Unlike a liquid fuel ramjet. and 40% AP. THE DUCTED ROCKET MOTOR An analytic model of a DR motor is based on Leingang & Petters [5]. While analyzing the results. Flight Mach Number for the DR The specific impulse is much greater than that of a solid rocket motor. in a solid fuel ducted rocket we can inject the fuel with high momentum from the gas generator. This is an addition to the model from [5]. an initial altitude of 47 kft (14325 m) provides enough air density for the DR burn. Therefore. starting the gravity turn with the 2nd stage. dx = V cos γ dt dh = V sin γ dt (8) (9) ⎛ 2 ⎞ dV mx ⎟ sin γ m mg =T − D−⎜ − ⎜ (R E + h ) ⎟ dt ⎝ ⎠ ⎛ 2 ⎞ dγ mx ⎟ cos γ mV mg = −⎜ − ⎜ (R E + h ) ⎟ dt ⎝ ⎠ T fuel mass = = m g 0 I SP burning time F = P2 a ( A4 − A2 a − AF ) = (10) = PT* A4 Z 4 − PT 2 a A2 a Z 2 a − PTF AF Z F The momentum function Z is defined by: (13) (11) Z= 1+ γ ⋅ M 2 ⎛ γ −1 ⎞ ⋅M 2⎟ ⎜1 + 2 ⎝ ⎠ γ γ −1 (14) (12) By using a Gravity turn trajectory.5 km. After burnout of the DR stage a pitch-up maneuver is performed to 34. this is not a full The DR characteristics are depicted in the following figures 5-7. a more adapted dimensional configuration can be developed. The calculations generally match the schematics in figure 4. an instant pitch up is assumed. The DR motor was designed to be placed behind a STAR48V 2nd stage motor. yet avoiding the damage from high dynamic pressure that would be encountered at lower altitude. 10% HTPB. it was evident that the dynamic pressure limits the performance of the 2nd stage. instead of launching at a level flight. we have a flight with zero angle of attack as a constraint that we utilize as an advantage. Figure 5. so its intakes are protruding on four corners circling the circumference of the 2nd stage. In this work the DR motor is used up to Mach 4. However. 4 AIAA/6th Responsive Space Conference 2008 . and embedded into this work for the 1st stage motor. Thrust Coefficient CT vs.2°. Release at 472 m/s (Mach 1. that will diminish the diameters of the 1st and 2nd motors (thus not using the COTS STAR48). The model was programmed in MATLAB and an investigation of the sensitivity to major parameters was done.5. A boron containing fuel has been used: 50% boron. The thrust coefficient in figure 5 increases until the point where the area ratio A∞/Ac reaches its maximum value of 1 and remains at that value. For the purpose of this work.AIAA-RS6-2008-5003 Equations of Motion The governing equations of motion for gravity turn trajectory are: design effort and if needed.6) ensures the DR ignition and operation. thus improving the motor's thrust by producing higher thrust coefficient for lower Mach numbers. The launch graphs are presented in figures 8-11.5 before pitching up for the 2nd stage. The 3rd motor kicks in for the acceleration into orbital velocity.32 m^2 divided between four identical intakes around the STAR48 2nd stage. the satellite is already at (or very close to) the required altitude. Figure 7. 5 AIAA/6th Responsive Space Conference 2008 . Several mechanisms for that pitch up are currently under investigation. Isp vs.AIAA-RS6-2008-5003 Figure 6. followed by a gravity turn trajectory till orbit insertion at 250 km altitude. In this launcher design. The downrange distance is 1287 km and the whole sequence lasts 365 seconds. Flight Mach Number for the DR The thrust is affected by the thrust coefficient and the increasing dynamic pressure as seen in figure 7. Figure 8.6 to Mach 4. Altitude vs. the air intake cross section A2 is 0. the launcher hardly loose velocity. and needs only acceleration to orbital velocity. Following 2nd stage is a 180 second coasting phase to orbit. Altitude vs. Figure 8 shows the constant flight path angle ascent of the 1st stage. DR Thrust vs. It can be seen in figure 10 that during the coasting phase between stages 2 & 3. Downrange Distance Figure 9 shows the assumption of an instant flight path angle change (pitch up) after the DR 1st stage burnout. The trajectory is optimized by the initial conditions. Figure 9. Gamma We can see (figure 10) the operation of the DR motor in accelerating the launcher from Mach 1. so that once the 3rd stage is initiated. Flight Mach Number RESULTS Calculations reveal that the resulting launcher is of 3085 kg mass and it can insert a 75 kg microsatellite into a 250X532 km orbit. 2° (+0.5. its burn time.2°.2° result in a lower orbit due to less than optimal DR stage contribution.AIAA-RS6-2008-5003 flight path angle. Velocity The DR motor is designed to work until its contribution is minimal (close to Mach 4. In analyzing various conditions. Sensitivity Analyses Satellite Mass & γ0 75 kg. The total launcher's mass is still 3085 kg. When adjusting the initial climb angle during the DR operation.8° (-0. the gravity turn trajectory's initial *By reducing the initial flight path angle by 0. 16° (-1kg) 76 kg. 16° (+1kg) 75 kg. However. the resulting trends presented hereafter are clear. This forces us to diminish the design burn time to 61 seconds (from the original 65). the apogee altitude is very sensitive to various launch parameters as will be presented hereafter. where the velocity curve is angling to horizontal at the end of its operation (can be seen in figure 11).2° initial flight path angle γ0 change. Time Since the residual velocity at burnout is 80.2°) for compensation. 16. the aim is to use the DR motor to increase the velocity up to the maximum of Mach 4.4) for the same amount of fuel.5. we receive an elliptic orbit of 250 X 532 km. Altitude vs. Velocity vs. the DR motor has higher airflow due to higher air density. thus affecting both the dynamic pressure and the motor performance. In this launch concept. the fuel should burn faster to maintain the same fuel/air ratio. the airflow density profile is changed due to density change during ascent. and the 2nd coast duration for inserting the satellite into the target orbit. 15. In order to elevate the perigee to 250 km we need to adjust the pitch up angle to 34. we receive the following results (Table 1): Table 1: Orbit [km] 250 X 532 251 X 596 249 X 469 *243 X 464 **240 X 429 Figure 10.5). This also may affect slightly the total launcher's mass.2°) Figure 11. Performing sensitivity analysis is more complex than with an all solid rocket motors launcher. the perigee is lower than the required 250 km.2°) 75 kg. 16° 74 kg. Some tradeoffs can be made between the 2nd coast duration and the final flight path angle (accepting a negative angle for an increased altitude and adjusting during orbital revolutions). This requires us to compensate with adjusting the DR fuel amount. The result is a shorter duration to reach the DR maximum velocity of Mach ~4. The steeper flight path angle prevents reaching the maximum required velocity (it reaches just Mach 4. This is a result of the CT behavior of the DR motor. since the DR performance is not constant and is dependent on the atmospheric conditions and on flight Mach number during its operation. Since the initial flight path angle is lower. SENSITIVITY ANALYSIS Several parameters were analyzed for the sensitivity to small changes. **using the same launcher configuration but increasing the initial flight path angle by 0. when taking a baseline case of a 75 kg satellite and analyzing its orbit sensitivity to a ±1 kg change or a ±0.55° (from the original 34. we receive a bit heavier launcher (3095 kg) and can reach a 250X526 km orbit. This results in a 250 X 441 km orbit. For example.5 m/s. By increasing the amount of fuel in the DR stage by 10 kg and thus its burn time to 69 seconds (from 65). 6 AIAA/6th Responsive Space Conference 2008 . In turn. we have to change the amount of fuel and burn time of the DR motor in order to support the target Mach number. Figure 13 presents the behavior of the required total launcher's mass for inserting a 75 kg microsatellite into a 250 X 532 km orbit per the required initial flight path angle. Therefore. without decreasing much the payload's initial mass. when increasing the initial velocity. However. Since a certain minimum velocity is required to ignite the DR motor. hence the dynamic pressure. 500 and 520 m/s) to show the trend. This will require a small increase in the DR fuel mass and an increase in the initial flight path angle. Since the initial γ0 is larger. by increasing the initial launch velocity. and insert a little lighter satellite into a much more elliptic orbit. γ0 By increasing the initial launch velocity we would expect higher performance. the launcher reaches the maximum DR 4. Figure 12. by adjusting some of the launch parameters.AIAA-RS6-2008-5003 Initial velocity sensitivity The sensitivity to a change in initial launch velocity was also analyzed.834 Total mass [kg] 3085 3050 3036 Orbit [km] 250X532 230X334 219X221 Figure 12 reveals the relation between the parameters. Initial Velocity Sensitivity Analyses DR burn [sec] 65 42 34 DR burnout alt [km] 29.095 23. the DR burn time is increased. assuming the pitch up remains the same and the total launcher's mass is almost unchanged. γ0 Tradeoffs can be made also between initial γ0 and the pitch up maneuver. However.5 Mach number sooner. This understanding allows us to tailor the orbit to our needs by placing the perigee above the area of interest. The second coast time may also diminish to decrees the deceleration during coasting ascent. the smaller the initial flight path angle can be for the same launcher. One can see that we can trade payload mass with additional ΔV at apogee burnout. Thus the main tradeoff is between the altitude and the initial flight path angle.551 21. Thus. Payload mass sensitivity The sensitivity of orbital performance to the payload mass was presented in table 1 above. A more detailed analysis is presented in figure 14. Initial Altitude Vs. The pitch up after DR Figure 13. the DR thrust and launcher's drag for altitudes of 14325. It is not necessarily intended for global operation. the higher we release the launcher. the asset in orbit will gain prolonged life. Table 2: V0 [m/s] 472 500 520 burnout will remain the same in this case. after burning less fuel and most importantly – at a lower altitude. we are subjected to atmospheric conditions. Using this method. The change will be in the burning time of the DR motor. The results are not optimized but present the trend and scale.5. Total Mass & Initial Velocity vs. When optimizing launch conditions with carrier aircraft performance envelope we'll receive the launch parameters definition per the mission requirements. three velocities were chosen (472. This point is important because it shows that one can use a standard launcher configuration even with an air breathing solid ducted rocket. The purpose of a tactical microsatellite is to operate above a specific area of interest. and allowing high 7 AIAA/6th Responsive Space Conference 2008 . Table 2 shows the effect on orbit characteristics. If we want to reach the same orbit for lower initial launch altitudes. like the pitch up angle. we have to incorporate an increase in the initial flight path angle to make sure we reach the same required orbit altitude of 250X532 km. we need to place the launcher at the same DR burnout altitude at its maximum velocity of Mach 4. 13000 and 11500 m were analyzed to show the trend. when employing a DR motor. Optimizing the parameters may result in changing other parameters that we kept constant for this analysis. Initial altitude sensitivity The initial release altitude affects the air density. A second use is the launch of a microsatellite into an elliptic orbit for an operation above a certain point (under the low perigee). possibly short lived satellites. into LEO on demand to mitigate a tactical need to replenish a loss of a strategic asset (a large higher altitude satellite). depending on the target orbit and allowable tradeoffs as presented in the above figures. The use of an air breathing DR motor for the first stage shows promising results and should be taken into consideration in developing tactical micro satellite launch vehicles. in the GTLV. This concept will enable the use of low-cost. CONCLUSIONS The use of an F15 as a platform for air launching of a tactical microsatellite via a small 3-stage combined airbreathing/solid rocket launcher has been demonstrated. The concept is proved to be viable. Figure 15. 8 AIAA/6th Responsive Space Conference 2008 . current results are promising. one can tailor efficiently the orbit in terms of inclination and time over target. GTLV orbit sensitivity: Apogee Altitude & ΔV at Perigee vs. Figure 15 shows the orbit performance sensitivity to initial flight path angle and launch altitude when inserted into a 250 km perigee. and the use of COTS motors decreases its cost. Comparison to an all-rocket launcher Another configuration of an all-solid rocket motor GTLV launcher is presented in [3] for a similar purpose of launching a microsatellite from a combat aircraft. In addition. Changes in other parameters (like mass and altitude) can often be remedied via adjustments in these two initial angles. In that concept. In figure 15 one can view as well the tradeoffs between payload mass and the gain of a more elliptic orbit per initial launch altitude. One is for launching a tactical reconnaissance microsatellite for a dedicated mission. thus prolonging its life for a couple of years at a fraction of the cost of launching a new satellite. a higher initial flight path angle was required but without angle changes during the launch sequence. It can be seen that even though a much heavier launcher (3900 kg) is used. a lighter payload can be inserted into similar orbits when compared to a combined air breathing/rocket launcher that is discussed in this work. while maintaining a high apogee for the required lifespan. γ0 was adjusted in the range of 62. Apogee Altitude & Perigee ΔV vs. A third use is for launching a guided motor to rendezvous with a decaying satellite. Satellite Mass Initial flight path angle sensitivity There are two initial flight path angles to work with – the initial release flight path angle and the post DR initial flight path angle for the gravity turn trajectory. for example. the launch was initiated directly into a gravity turn trajectory (since there was no air breathing engine).5°-67.2° for inserting the satellite into the same 250 km perigee. Figure 14. and flight envelope.AIAA-RS6-2008-5003 apogee for longevity. Satellite Mass Even though the combined air-breathing/solid rocket launcher solution is not yet optimized. POTENTIAL USE There can be several uses for this concept of launching a small payload to LEO. Because of the air launching. development complexity and carrier aircraft adaptability in terms of structural modification. VA. J. Vol. "Air Launch Trajectories to Earth Orbit". J. N. 6. J.G. 6_2703 . W. Vol 59. May 10-14. 1996. Progress in Astronautics and Aeronautics. 5. 2003. and Petters.12 July 2006. 2003. A. J. 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