Investigation of Combined Air-BreathingRocket Propulsion for Air Launch of Micro

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6th Responsive Space ConferenceApril 28May 1, 2008

Los Angeles, CA

Investigation of Combined Air-

breathing/Rocket Propulsion forAir Launch of Micro-Satellitesfrom a Combat Aircraft

Avichai Socher and Alon Gany

Faculty of Aerospace EngineeringTechnion - Israel Institute of TechnologyHaifa 32000, Israel

6th Responsive Space Conference

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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 32000, Israel; 972-3-5325283

[email protected]

Alon Gany

Technion Israel Institute of Technology

Faculty of Aerospace Engineering, Haifa 32000, Israel; 972-4-8292554

[email protected]

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 ground-based 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) 1st 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 flightpath angle, coasting time, and ducted rocket sizing was done. 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. 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^2]

A2 Intake area [m^2]

A2a Air entrance to combustion chamber area

[m^2]A4 Post combustion chamber area [m^2]

A Free stream intake area [m^2]

AC Gas generator burn area [m^2]AF Gas generator exit area [m^2]B Ballistic coefficient [kg/m^2]

CD Drag coefficient

COTS Commercial Off The ShelfCT4 Thrust coefficient at engine station 4

D Drag [N]

DR Ducted Rocket

F Of the fuelg Gravity [m/s^2]

GTLV Gravity Turn Launch Vehicle

h Altitude [km]

LEO Low Earth Orbit

m Mass [kg]

MSLV Micro Satellite Launch VehicleORS Operational Responsive Space

P Pressure

r Radius from earth's centerRE Earth's radius [km]

T Thrust [N]

USAF United States Air ForceV Velocity [m/s]

Vc Circular orbit velocity

X Down range distance [km]

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Z Momentum function

r Change in orbit radius Density [kg/m^3]

m Fuel consumption rate [kg/sec] Period [sec]

Flight path angle [deg]

GME [m^3/sec^2]

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 highsensitivity 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. Dueto 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 0.25 m^2 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, 290 and 250 km.

The satellite's lifespan is dependent on the ballistic

coefficient of the satellite B:

=

2m

kg

AC

mB

D

(1)

A typical value for CD is 2.2.

The change in orbit's radius is defined as follows:

=

rrevolution

r (2)

= r

Br

(3)

=

3

2r

(4)

Incorporating Eqs. (3) and (4) into Eq. (2) results in:

Br

revolution

r

22=

(5)

The additional speed required to return the satellitefrom its current position to its original orbit is obtained

from total energy calculations via a Hohmann

maneuver.

The amount of fuel required to gain the said V is

defined by:

=

SP

i

Ig

V

ifemm 1 (6)

It can be seen in figure 1 that at a 290 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 acloser-to-horizontal launch, concluding that launching

from high altitudes can significantly reduce the

practical size of launch vehicles, especially if a highacceleration is associated with the selected propulsion

technology. This work complies with his findings andrecommendations for air breathing small launchers.

Figure 1. Orbit Decay vs. Time

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).

Illustration of the launcher's preliminary configurationis presented in figure 2.

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Figure 2. 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 microsatellitefrom a combat aircraft focused on all rocket, 3-stagelaunchers. Among them is the F15 MSLV with a 4500

kg, 6.7m long launcher that could insert a 93 kg

payload into a circular 225 km orbit [1], [2]. A similar

launcher that operates entirely via a Gravity Turn

Trajectory is the 3900 kg, 6.2m long Gravity Turn

Launch Vehicle (GTLV) launcher that could insert a 75kg payload into a 250 km orbit from similar initial

conditions [3].

Figure 3. Pre-Launch Launcher's Mounting

Savu [4] analyzed the launching of an 800 kg rocketwith a 10 kg nano-satellite as payload, from a MiG-21

military aircraft into a 116 km orbit.

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-third-

size Pegasus XL, the F-5F Tiger II with one-half-size

Pegasus XL, and the F-4E Phantom II with two-thirds-size Pegasus XL. The payloads were 36, 122 and 289

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 1st stage

via an air-breathing motor.

Estimated Specific impulse (Isp) as a function of flightMach 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 proposedin [1] & [3] is not applicable since the launcher will

pass through the atmosphere too fast and will enter atoo 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 air-

breathing 1st stage.

Figure 4. Isp vs. Mach Number for different

Engines

In order to simplify the solution, the use of a Gravity

Turn Trajectory is proposed for the post-DR launch

sequence.

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Equations of Motion

The governing equations of motion for gravity turn

trajectory are:

cosVdt

dx= (8)

sinVdt

dh= (9)

( )sin

2

+=

hR

xmmgDT

dt

dVm

E

(10)

( )

cos

2

+=

hR

xmmg

dt

dmV

E

(11)

timeburning

massfuel

Ig

TmSP 0

== (12)

By using a Gravity turn trajectory, we have a flight withzero angle of attack as a constraint that we utilize as an

advantage.

The model was programmed in MATLAB and aninvestigation of the sensitivity to major parameters was

done.

While analyzing the results, it was evident that thedynamic pressure limits the performance of the 2nd

stage. Therefore, instead of launching at a level flight, amoderate 16 flight path angle was chosen for the DRmotor stage. Release at 472 m/s (Mach 1.6) ensures the

DR ignition and operation; an initial altitude of 47 kft

(14325 m) provides enough air density for the DR burn,

yet avoiding the damage from high dynamic pressure

that would be encountered at lower altitude.

After burnout of the DR stage a pitch-up maneuver is

performed to 34.2, starting the gravity turn with the 2nd

stage. For the purpose of this work, an instant pitch upis 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 1ststage motor. The DR motor was designed to be placed

behind a STAR48V 2nd stage motor, so its intakes are

protruding on four corners circling the circumference of

the 2nd stage. It is assumed that the configuration can beinstalled under the F15 belly. However, this is not a full

design effort and if needed, a more adapted dimensional

configuration can be developed, that will diminish the

diameters of the 1st and 2nd 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 byproducing higher thrust coefficient for lower Mach

numbers. This is an addition to the model from [5].

( )

FFTFaaaTT

Faa

ZAPZAPZAP

AAAPF

=

==

22244

*

242(13)

The momentum function Z is defined by:

12

2

2

11

1

+

+=

M

MZ (14)

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 theDR motor is used up to Mach 4.5.

Figure 5. Thrust Coefficient CT vs. Flight Mach

Number for the DR

The specific impulse is much greater than that of a solid

rocket motor. The calculations generally match theschematics in figure 4.

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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 A2 is 0.32

m^2 divided between four identical intakes around theSTAR48 2nd stage.

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 microsatelliteinto a 250X532 km orbit.

The launch graphs are presented in figures 8-11.

Figure 8 shows the constant flight path angle ascent of

the 1st stage, followed by a gravity turn trajectory till

orbit insertion at 250 km altitude. The downrange

distance is 1287 km and the whole sequence lasts 365

seconds.

Figure 8. Altitude vs. Downrange Distance

Figure 9 shows the assumption of an instant flight path

angle change (pitch up) after the DR 1st stage burnout.

Several mechanisms for that pitch up are currently

under investigation. It can be seen in figure 10 thatduring the coasting phase between stages 2 & 3, the

launcher hardly loose velocity. The trajectory is

optimized by the initial conditions, so that once the 3 rdstage is initiated, the satellite is already at (or very close

to) the required altitude, and needs only acceleration to

orbital velocity.

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 4.5before pitching up for the 2

nd stage. Following 2nd stage

is a 180 second coasting phase to orbit. The 3rd motor

kicks in for the acceleration into orbital velocity.

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Figure 10. Altitude vs. Velocity

The DR motor is designed to work until its contribution

is minimal (close to Mach 4.5), where the velocitycurve is angling to horizontal at the end of its operation

(can be seen in figure 11). This is a result of the CTbehavior of the DR motor.

Figure 11. Velocity vs. Time

Since the residual velocity at burnout is 80.5 m/s, wereceive an elliptic orbit of 250 X 532 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 motorto increase the velocity up to the maximum of Mach

4.5. In analyzing various conditions, we have to changethe amount of fuel and burn time of the DR motor inorder 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. Thisrequires us to compensate with adjusting the DR fuel

amount, its burn time, the gravity turn trajectory's initial

flight path angle, and the 2nd 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 2nd coast duration and the final flight

path angle (accepting a negative angle for an increasedaltitude and adjusting during orbital revolutions).

Performing sensitivity analysis is more complex thanwith 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.

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 orbitsensitivity to a 1 kg change or a 0.2 initial flight

path angle 0 change, we receive the following results(Table 1):

Table 1: Sensitivity Analyses

Orbit [km] Satellite Mass & 0

250 X 532 75 kg, 16

251 X 596 74 kg, 16 (-1kg)

249 X 469 76 kg, 16 (+1kg)

*243 X 464 75 kg, 15.8 (-0.2)

**240 X 429 75 kg, 16.2 (+0.2)

*By reducing the initial flight path angle by 0.2, 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 reachthe DR maximum velocity of Mach ~4.5. This forces us

to diminish the design burn time to 61 seconds (fromthe 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 250 km. In order to

elevate the perigee to 250 km we need to adjust the

pitch up angle to 34.55 (from the original 34.2) for

compensation. This results in a 250 X 441 km orbit.

**using the same launcher configuration but increasing

the initial flight path angle by 0.2 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 250X526 km orbit.

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Initial velocity sensitivity

The sensitivity to a change in initial launch velocity

was also analyzed. Since a certain minimum velocity isrequired to ignite the DR motor, three velocities were

chosen (472, 500 and 520 m/s) to show the trend. Table

2 shows the effect on orbit characteristics.

Table 2: Initial Velocity Sensitivity Analyses

V0

[m/s]

Orbit

[km]

DR burn

[sec]

DR burnout

alt [km]

Total mass

[kg]

472 250X532 65 29.095 3085

500 230X334 42 23.551 3050

520 219X221 34 21.834 3036

Figure 12 reveals the relation between the parameters.

Figure 12. Total Mass & Initial Velocity vs. 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 4.5Mach number sooner, after burning less fuel and most

importantly at a lower altitude. Therefore, when

increasing the initial velocity, we have to incorporate anincrease in the initial flight path angle to make sure we

reach the same required orbit altitude of 250X532 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 dragfor altitudes of 14325, 13000 and 11500 m were

analyzed to show the trend. If we want to reach the

same orbit for lower initial launch altitudes, we need toplace 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

burnout will remain the same in this case. Since the

initial 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 250 X 532 km orbit per the required initial flight pathangle, 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 parametersmay result in changing other parameters that we kept

constant for this analysis, like the pitch up angle.

Figure 13. Initial Altitude Vs. 0

Tradeoffs can be made also between initial 0 and the

pitch up maneuver. However, the higher we release thelauncher, the smaller the initial flight path angle can be

for the same launcher. The change will be in theburning time of the DR motor. This point is important

because it shows that one can use a standard launcherconfiguration 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 V at apogee burnout, and insert a little

lighter satellite into a much more elliptic orbit. Usingthis 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 necessarilyintended 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

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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 V 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 twoinitial 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 motorGTLV 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 agravity turn trajectory (since there was no air breathing

engine). Figure 15 shows the orbit performance

sensitivity to initial flight path angle and launch altitudewhen inserted into a 250 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, inthe GTLV, a higher initial flight path angle was

required but without angle changes during the launchsequence. In figure 15 one can view as well the

tradeoffs between payload mass and the gain of a moreelliptic orbit per initial launch altitude. 0 was adjusted

in the range of 62.5-67.2 for inserting the satellite into

the same 250 km perigee.

Figure 15. GTLV orbit sensitivity: Apogee Altitude

& V at Perigee vs. Satellite Mass

Even though the combined air-breathing/solid rocketlauncher 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 tacticalreconnaissance 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 highapogee for the required lifespan.

A third use is for launching a guided motor torendezvous 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 air-

breathing/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 ofstructural 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.

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REFERENCES

1. Hague, N., Siegenthaler, E., and Rothman, J.,Enabling Responsive Space: F-15 MicrosatelliteLaunch Vehicle, Proceedings of the Aerospace

Conference, 2003. 2003 IEEE Volume 6, March

8-15, 2003, pp. 6_2703 - 6_2708, IEEEAC paper

#1102.

2. Rothman, J. and Siegenthaler, E., "ResponsiveSpace Launch - The F-15 Microsatellite Launch

Vehicle", AIAA-LA Section/SSTC 2003-9002,

1st Responsive Space Conference, April 13,

2003, Redondo Beach, CA.

3. Socher, A. and Gany, A., "A ParametricInvestigation of a Propulsion System for Air

Launch of Micro-Satellites from a Combat

Aircraft", The 48th Israel Annual Conference on

Aerospace Sciences, February 27-28, 2008, Tel

Aviv & Haifa, Israel.

4. Savu, G., "Micro, Nano and Pico satellitesLaunched from the Romanian Territory," ActaAstronautica, Vol 59, 2006, pp. 858 861.

5. Leingang, J. L. and Petters, D. P., "DuctedRockets", in: Tactical missile propulsion,

Progress in Astronautics and Aeronautics. Vol.

170, AIAA, Inc, Reston, VA, 1996, pp. 447-468.

6. Boltz, F. W.,"Low-Cost Small-Satellite DeliverySystem," Journal of Spacecraft and Rockets, Vol.

39, No. 5, 2002, pp. 818820.

7. Segal, C., "Propulsion Systems for HypersonicFlight", in Fundamentals of Hypersonic Flow -

Aerothermodynamics, Critical Technologies forHypersonic Vehicle Development, (D.G. Fletcher,

ed.), VKI RP 2004-21, May 10-14, 2004, Rhode

Saint Gense, Belgium.

8. Whitehead, J. C., "Air Launch Trajectories toEarth Orbit", 42nd AIAA/ASME/SAE/ASEE Joint

Propulsion Conference & Exhibit, 9 - 12 July

2006, Sacramento, California, AIAA 2006-4785.

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