Investigation of Combined Airbreathing/Rocket. Air Launch of Micro-Satellites from a Combat Aircraft

Transcription

1 6th Responsive Space Conference AIAA-RS Investigation 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 3000, Israel 6th Responsive Space Conference April 8 May 1, 008 Los Angeles, CA

2 Investigation of Combined Air-breathing/Rocket Propulsion for Air Launch of Micro- Satellites from a Combat Aircraft Avichai Socher Technion Israel Institute of Technology Faculty of Aerospace Engineering, Haifa 3000, Israel; asocher@netvision.net.il Alon Gany Technion Israel Institute of Technology Faculty of Aerospace Engineering, Haifa 3000, Israel; gany@tx.technion.ac.il ABSTRACT This work presents the analytical results of a parametric investigation of a launch concept of micro-satellites from a combat aircraft. The concept of air launching of a satellite from a carrier aircraft is not new; however, most designs consider heavy aircraft and launch vehicle to place a mini to a large satellite, typically launched today via groundbased rocket launchers. Documented air launcher designs usually incorporate a lift aided trajectory. It is the authors intention to present a method for air launching of a low-cost tactical micro-satellite, on demand, for various missions, using a weight economical vehicle via a Gravity Turn Trajectory. The carrier aircraft will be an F-15 fighter, and the launcher will be a 3-stage vehicle, assembled from a ducted rocket (ramrocket) 1 st stage and two solid propellant rocket stages. The option of an air-breathing engine for the first stage results from the high initial speed (as high as Mach 1.6) provided to the launcher by the carrier aircraft. An air-breathing engine provides much higher energetic performance compared to a standard solid rocket motor (higher Isp and lower mass). A ducted rocket was chosen over other ramjet configurations for its higher thrust coefficient. Optimization on initial flight path angle, coasting time, and ducted rocket sizing was done. The solution presents a concept for placing a kg micro-satellite in either a circular 50 km low earth orbit (LEO) or a more elliptic LEO. It is demonstrated that an air-launch of a micro-satellite from a combat aircraft is a viable solution. KEYWORDS: Air Launch, Microsatellite, Ducted Rocket, Ramjet NOMENCLATURE A Cross section of a satellite [m^] A Intake area [m^] A a Air entrance to combustion chamber area [m^] A 4 Post combustion chamber area [m^] A Free stream intake area [m^] A C Gas generator burn area [m^] A F Gas generator exit area [m^] B Ballistic coefficient [kg/m^] C D Drag coefficient COTS Commercial Off The Shelf C T4 Thrust coefficient at engine station 4 D Drag [N] DR Ducted Rocket F Of the fuel g Gravity [m/s^] GTLV Gravity Turn Launch Vehicle h Altitude [km] LEO Low Earth Orbit m Mass [kg] MSLV Micro Satellite Launch Vehicle ORS Operational Responsive Space P Pressure r Radius from earth's center R E Earth's radius [km] T Thrust [N] USAF United States Air Force V Velocity [m/s] Vc Circular orbit velocity X Down range distance [km] 1

3 Z Momentum function Δr Change in orbit radius ρ Density [kg/m^3] m Fuel consumption rate [kg/sec] τ Period [sec] γ Flight path angle [deg] µ GM E [m^3/sec^] INTRODUCTION Investigating the sensitivity of the lifespan to the satellite's cross section and initial orbit altitude, one learns that for low orbit altitudes, there is high sensitivity to dimensions and mass. 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. Due to those payload characteristics, it can be launched via air launch from a combat aircraft, thus creating a tactical responsive capability. For a microsatellite with a 0.5 m^ cross section (like a 0.5m cube), mass of 75 kg and fuel amount of 10 kg we receive the following results shown in Figure 1 for two initial altitudes, 90 and 50 km. from total energy calculations via a Hohmann maneuver. The amount of fuel required to gain the said ΔV is defined by: ΔV i Δ = g I SP m f mi 1 e (6) It can be seen in figure 1 that at a 90 km circular orbit, a once-a-month altitude motor boost is required to enable a 306-day operation. The red diamond in figure 1 marks the fuel end point. Whitehead [8] also showed the benefits of air launching via a small launcher into LEO, and the efficiency of a closer-to-horizontal launch, concluding that launching from high altitudes can significantly reduce the practical size of launch vehicles, especially if a high acceleration is associated with the selected propulsion technology. This work complies with his findings and recommendations for air breathing small launchers. The satellite's lifespan is dependent on the ballistic coefficient of the satellite B: B m kg C A m = D (1) A typical value for C D is.. The change in orbit's radius is defined as follows: Δr = r τ () revolution ρ r = μ r (3) B τ = π 3 r ϖ Incorporating Eqs. (3) and (4) into Eq. () results in: Δr ρ = πr revolution B (4) (5) The additional speed required to return the satellite from its current position to its original orbit is obtained Figure 1. Orbit Decay vs. Time CONCEPT AND PRELIMINARY DESIGN The concept presented in this work is a 3-stage launcher consisting of a 1 st stage of ducted rocket (DR) motor (also called ramrocket) and two solid rocket COTS motors (STAR48V and STAR7). Illustration of the launcher's preliminary configuration is presented in figure.

4 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, the F-5F Tiger II with one-half-size Pegasus XL, and the F-4E Phantom II with two-thirdssize Pegasus XL. The payloads were 36, 1 and 89 lb, respectively. What unifies all those concepts is the fact that they all use solid rocket motors. By that, the launcher carries all the propellant and oxidizer on board. Since the first stage of the launch passes through the atmosphere, we can use the oxygen in the atmosphere for the 1 st stage via an air-breathing motor. Figure. Launcher Schematic Configuration The launcher is carried under the belly of an F15 fighter aircraft as shown in figure 3. Past investigations of air launching of a microsatellite from a combat aircraft focused on all rocket, 3-stage launchers. Among them is the F15 MSLV with a 4500 kg, 6.7m long launcher that could insert a 93 kg payload into a circular 5 km orbit [1], []. A similar launcher that operates entirely via a Gravity Turn Trajectory is the 3900 kg, 6.m long Gravity Turn Launch Vehicle (GTLV) launcher that could insert a 75 kg payload into a 50 km orbit from similar initial conditions [3]. 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]. 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. Therefore, a level flight or a moderate ascent is required, when using an airbreathing 1 st stage. Figure 4. Isp vs. Mach Number for different Engines Figure 3. Pre-Launch Launcher's Mounting Savu [4] analyzed the launching of an 800 kg rocket with a 10 kg nano-satellite as payload, from a MiG-1 military aircraft into a 116 km orbit. In order to simplify the solution, the use of a Gravity Turn Trajectory is proposed for the post-dr launch sequence. 3

5 Equations of Motion The governing equations of motion for gravity turn trajectory are: dx = V cosγ (8) dt dh = V sin γ (9) dt dv m dt = T D mg dγ mv = mg dt mx mx ( R + h) E ( R + h) E cosγ sinγ (10) (11) T fuel mass m = = (1) g I burning time 0 SP By using a Gravity turn trajectory, we have a flight with zero angle of attack as a constraint that we utilize as an advantage. The model was programmed in MATLAB and an investigation of the sensitivity to major parameters was done. While analyzing the results, it was evident that the dynamic pressure limits the performance of the nd stage. Therefore, instead of launching at a level flight, a moderate 16 flight path angle was chosen for the DR motor stage. Release at 47 m/s (Mach 1.6) ensures the DR ignition and operation; an initial altitude of 47 kft (1435 m) provides enough air density for the DR burn, yet avoiding the damage from high dynamic pressure that would be encountered at lower altitude. design effort and if needed, a more adapted dimensional configuration can be developed, that will diminish the diameters of the 1 st and nd motors (thus not using the COTS STAR48). Unlike a liquid fuel ramjet, in a solid fuel ducted rocket we can inject the fuel with high momentum from the gas generator, thus improving the motor's thrust by producing higher thrust coefficient for lower Mach numbers. This is an addition to the model from [5]. F = P = P a * T A4 ( A A A ) Z 4 4 P a T a A a Z F a = P The momentum function Z is defined by: TF A F Z F (13) Z = 1+ γ M γ (14) γ 1 1 γ 1+ M The DR characteristics are depicted in the following figures 5-7. A boron containing fuel has been used: 50% boron, 10% HTPB, and 40% AP. The following figures 5-7 are for a constant altitude of 11.5 km. 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. In this work the DR motor is used up to Mach 4.5. After burnout of the DR stage a pitch-up maneuver is performed to 34., starting the gravity turn with the nd stage. For the purpose of this work, an instant pitch up is assumed. THE DUCTED ROCKET MOTOR An analytic model of a DR motor is based on Leingang & Petters [5], and embedded into this work for the 1 st stage motor. The DR motor was designed to be placed behind a STAR48V nd stage motor, so its intakes are protruding on four corners circling the circumference of the nd stage. It is assumed that the configuration can be installed under the F15 belly. However, this is not a full Figure 5. Thrust Coefficient C T vs. Flight Mach Number for the DR The specific impulse is much greater than that of a solid rocket motor. The calculations generally match the schematics in figure 4. 4

6 Figure 6. Isp vs. Flight Mach Number for the DR The thrust is affected by the thrust coefficient and the increasing dynamic pressure as seen in figure 7. In this launcher design, the air intake cross section A is 0.3 m^ divided between four identical intakes around the STAR48 nd stage. Figure 8. Altitude vs. Downrange Distance Figure 9 shows the assumption of an instant flight path angle change (pitch up) after the DR 1 st stage burnout. Several mechanisms for that pitch up are currently under investigation. It can be seen in figure 10 that during the coasting phase between stages & 3, the launcher hardly loose velocity. The trajectory is optimized by the initial conditions, so that once the 3 rd stage is initiated, the satellite is already at (or very close to) the required altitude, and needs only acceleration to orbital velocity. Figure 7. DR Thrust vs. 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 50X53 km orbit. The launch graphs are presented in figures Figure 8 shows the constant flight path angle ascent of the 1 st stage, followed by a gravity turn trajectory till orbit insertion at 50 km altitude. The downrange distance is 187 km and the whole sequence lasts 365 seconds. Figure 9. Altitude vs. Gamma We can see (figure 10) the operation of the DR motor in accelerating the launcher from Mach 1.6 to Mach 4.5 before pitching up for the nd stage. Following nd stage is a 180 second coasting phase to orbit. The 3 rd motor kicks in for the acceleration into orbital velocity. 5

7 flight path angle, and the nd coast duration for inserting the satellite into the target orbit. This also may affect slightly the total launcher's mass. Some tradeoffs can be made between the nd coast duration and the final flight path angle (accepting a negative angle for an increased altitude and adjusting during orbital revolutions). Performing sensitivity analysis is more complex than with an all solid rocket motors launcher, since the DR performance is not constant and is dependent on the atmospheric conditions and on flight Mach number during its operation. However, the resulting trends presented hereafter are clear. Figure 10. Altitude vs. Velocity The DR motor is designed to work until its contribution is minimal (close to Mach 4.5), where the velocity curve is angling to horizontal at the end of its operation (can be seen in figure 11). This is a result of the C T behavior of the DR motor. SENSITIVITY ANALYSIS Several parameters were analyzed for the sensitivity to small changes. For example, when taking a baseline case of a 75 kg satellite and analyzing its orbit sensitivity to a ±1 kg change or a ±0. initial flight path angle γ 0 change, we receive the following results (Table 1): Table 1: Sensitivity Analyses Figure 11. Velocity vs. Time Since the residual velocity at burnout is 80.5 m/s, we receive an elliptic orbit of 50 X 53 km. the apogee altitude is very sensitive to various launch parameters as will be presented hereafter. In this launch concept, the aim is to use the DR motor to increase the velocity up to the maximum of Mach 4.5. In analyzing various conditions, we have to change the amount of fuel and burn time of the DR motor in order to support the target Mach number. When adjusting the initial climb angle during the DR operation, the airflow density profile is changed due to density change during ascent, thus affecting both the dynamic pressure and the motor performance. This requires us to compensate with adjusting the DR fuel amount, its burn time, the gravity turn trajectory's initial Orbit [km] Satellite Mass & γ 0 50 X kg, X kg, 16 (-1kg) 49 X kg, 16 (+1kg) *43 X kg, 15.8 (-0. ) **40 X kg, 16. (+0. ) *By reducing the initial flight path angle by 0., the DR motor has higher airflow due to higher air density. In turn, the fuel should burn faster to maintain the same fuel/air ratio. The result is a shorter duration to reach the DR maximum velocity of Mach ~4.5. This forces us to diminish the design burn time to 61 seconds (from the original 65). The total launcher's mass is still 3085 kg. Since the initial flight path angle is lower, the perigee is lower than the required 50 km. In order to elevate the perigee to 50 km we need to adjust the pitch up angle to (from the original 34. ) for compensation. This results in a 50 X 441 km orbit. **using the same launcher configuration but increasing the initial flight path angle by 0. result in a lower orbit due to less than optimal DR stage contribution. The steeper flight path angle prevents reaching the maximum required velocity (it reaches just Mach 4.4) for the same amount of fuel. By increasing the amount of fuel in the DR stage by 10 kg and thus its burn time to 69 seconds (from 65), we receive a bit heavier launcher (3095 kg) and can reach a 50X56 km orbit. 6

8 Initial velocity sensitivity The sensitivity to a change in initial launch velocity was also analyzed. Since a certain minimum velocity is required to ignite the DR motor, three velocities were chosen (47, 500 and 50 m/s) to show the trend. Table shows the effect on orbit characteristics. V0 [m/s] Table : Orbit [km] Initial Velocity Sensitivity Analyses DR burn [sec] DR burnout alt [km] Total mass [kg] 47 50X X X Figure 1 reveals the relation between the parameters. burnout will remain the same in this case. Since the initial γ 0 is larger, the DR burn time is increased. The second coast time may also diminish to decrees the deceleration during coasting ascent. Figure 13 presents the behavior of the required total launcher's mass for inserting a 75 kg microsatellite into a 50 X 53 km orbit per the required initial flight path angle, assuming the pitch up remains the same and the total launcher's mass is almost unchanged. Thus the main tradeoff is between the altitude and the initial flight path angle. The results are not optimized but present the trend and scale. Optimizing the parameters may result in changing other parameters that we kept constant for this analysis, like the pitch up angle. Figure 1. Total Mass & Initial Velocity vs. γ 0 By increasing the initial launch velocity we would expect higher performance. However, when employing a DR motor, we are subjected to atmospheric conditions. Thus, by increasing the initial launch velocity, the launcher reaches the maximum DR 4.5 Mach number sooner, after burning less fuel and most importantly at a lower altitude. Therefore, when increasing the initial velocity, we have to incorporate an increase in the initial flight path angle to make sure we reach the same required orbit altitude of 50X53 km. When optimizing launch conditions with carrier aircraft performance envelope we'll receive the launch parameters definition per the mission requirements. Initial altitude sensitivity The initial release altitude affects the air density; hence the dynamic pressure, the DR thrust and launcher's drag for altitudes of 1435, and m were analyzed to show the trend. If we want to reach the same orbit for lower initial launch altitudes, we need to place the launcher at the same DR burnout altitude at its maximum velocity of Mach 4.5. This will require a small increase in the DR fuel mass and an increase in the initial flight path angle. The pitch up after DR Figure 13. Initial Altitude Vs. γ 0 Tradeoffs can be made also between initial γ 0 and the pitch up maneuver. However, the higher we release the launcher, the smaller the initial flight path angle can be for the same launcher. The change will be in the burning time of the DR motor. This point is important because it shows that one can use a standard launcher configuration even with an air breathing solid ducted rocket, by adjusting some of the launch parameters. 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. One can see that we can trade payload mass with additional ΔV at apogee burnout, and insert a little lighter satellite into a much more elliptic orbit. Using this method, the asset in orbit will gain prolonged life, without decreasing much the payload's initial mass. The purpose of a tactical microsatellite is to operate above a specific area of interest. It is not necessarily intended for global operation. This understanding allows us to tailor the orbit to our needs by placing the perigee above the area of interest, and allowing high 7

9 apogee for longevity. Because of the air launching, one can tailor efficiently the orbit in terms of inclination and time over target. Figure 14. Apogee Altitude & Perigee ΔV vs. 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. Changes in other parameters (like mass and altitude) can often be remedied via adjustments in these two initial angles, depending on the target orbit and allowable tradeoffs as presented in the above figures. 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 that concept, the launch was initiated directly into a gravity turn trajectory (since there was no air breathing engine). Figure 15 shows the orbit performance sensitivity to initial flight path angle and launch altitude when inserted into a 50 km perigee. 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. In addition, in the GTLV, a higher initial flight path angle was required but without angle changes during the launch sequence. 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. γ 0 was adjusted in the range of for inserting the satellite into the same 50 km perigee. Figure 15. GTLV orbit sensitivity: Apogee Altitude & ΔV at Perigee vs. Satellite Mass Even though the combined air-breathing/solid rocket launcher solution is not yet optimized, current results are promising. POTENTIAL USE There can be several uses for this concept of launching a small payload to LEO. One is for launching a tactical reconnaissance microsatellite for a dedicated mission. This concept will enable the use of low-cost, 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), for example. A second use is the launch of a microsatellite into an elliptic orbit for an operation above a certain point (under the low perigee), while maintaining a high apogee for the required lifespan. A third use is for launching a guided motor to rendezvous with a decaying satellite, thus prolonging its life for a couple of years at a fraction of the cost of launching a new satellite. 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, and the use of COTS motors decreases its cost, development complexity and carrier aircraft adaptability in terms of structural modification, and flight envelope. 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. 8

10 REFERENCES 1. Hague, N., Siegenthaler, E., and Rothman, J., Enabling Responsive Space: F-15 Microsatellite Launch Vehicle, Proceedings of the Aerospace Conference, IEEE Volume 6, March 8-15, 003, pp. 6_703-6_708, IEEEAC paper #110.. Rothman, J. and Siegenthaler, E., "Responsive Space Launch - The F-15 Microsatellite Launch Vehicle", AIAA-LA Section/SSTC , 1st Responsive Space Conference, April 1 3, 003, Redondo Beach, CA. 3. Socher, A. and Gany, A., "A Parametric Investigation of a Propulsion System for Air Launch of Micro-Satellites from a Combat Aircraft", The 48th Israel Annual Conference on Aerospace Sciences, February 7-8, 008, Tel Aviv & Haifa, Israel. 4. Savu, G., "Micro, Nano and Pico satellites Launched from the Romanian Territory," Acta Astronautica, Vol 59, 006, pp Leingang, J. L. and Petters, D. P., "Ducted Rockets", in: Tactical missile propulsion, Progress in Astronautics and Aeronautics. Vol. 170, AIAA, Inc, Reston, VA, 1996, pp Boltz, F. W.,"Low-Cost Small-Satellite Delivery System," Journal of Spacecraft and Rockets, Vol. 39, No. 5, 00, pp Segal, C., "Propulsion Systems for Hypersonic Flight", in Fundamentals of Hypersonic Flow - Aerothermodynamics, Critical Technologies for Hypersonic Vehicle Development, (D.G. Fletcher, ed.), VKI RP 004-1, May 10-14, 004, Rhode Saint Genèse, Belgium. 8. Whitehead, J. C., "Air Launch Trajectories to Earth Orbit", 4nd AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, 9-1 July 006, Sacramento, California, AIAA