1. Orbital Payload Delivery Using Hydrogen and Hydrocarbon Fuelled Scramjet Engines
M. R. Tetlow and C.J. Doolan
School on Mechanical Engineering
The University of Adelaide

2. Overview Current launch systems Scramjet background
Mission profile and vehicle description
Software operation
Trajectory outputs
Analysis of results
Conclusions

3. Aim
Design a mission using a hydrocarbon powered (JetA) and a hydrogen powered scramjet stage to reach a 200km circular orbit
Compare the mission profiles and performance of the two launch systems
Compare the performance to current rocket powered systems

4. Current Launch Systems
Launch Vehicle
Payload mass (mass fraction)
LEO orbit and inclination
ASLV
150kg (0.36%)
400km at 43°
M-3S11
780kg (1.26%)
185km at 31°
Long March CZ1D
720kg (0.9%)
200km at 28°
Start-1
360kg (0.6%)
400km at 90°
1% at 200km is indicative of the performance of this class of vehicle [Isakowitz ]

5. Scramjets Supersonic combustion ramjet Hydrogen fuelled
Geometry dependent on operating conditions
Hydrogen fuelled
High energy, low storage density
Operating range: Mach 5 to 15
Isp ~ 3000s
Hydrocarbon fuelled
Lower energy, high storage density
Operating range: Mach 5 to 10
Isp ~1200s
Minimum dynamic pressure ~10kPa

6. Waveriders

7. Waveriders Blended wing vehicle with integrated propulsion system
“Ride” the shock wave
Aerodynamics are Mach No. dependent
Fuel mass fractions
ε=0.58 for hydrogen fuelled vehicle
ε=0.7 for hydrocarbon fuelled vehicle
ε=0.9 for rockets

8. Quasi-1D Scramjet Propulsion Model
Flow From Inlet
Displacement Thickness Growth
Combustion
Area Change
Shear Stress
Ignition Delay
Heat Transfer
Injector

9. Quasi-1D Scramjet Propulsion Model
Set of ODEs used to describe scramjet propulsion.
2-step chemistry model.
Skin friction and wall heat transfer included.
H2 and Jet A fuel options.
Idealised hypersonic inlet (with losses) used to supply combustor.
Lawrence Livermore ODEPACK Solver used for ODE solution.

10. T4 Experiment
Parallel Combustor
Scramjet model validated against shock tunnel data (T4, University of Queensland, Boyce et al., 2000).
Parallel and diverging combustor data used for validation study.
Good agreement obtained using an 88% combustion efficiency.
A conservative 50% combustion efficiency was used for trajectory modelling (for combustor losses).
Diverging Combustor

11. Common Design Parameters
GLOW 9300kg
2 stage solid rocket booster
Stage 1: 2420kg start mass, 1980kg propellant
Stage 2: 4880kg start mass, 4000kg propellant
Cranked wing concept with aerodynamics taken from a NASA study
Rocket powered upper stage with performance based on the H2 upper stage

12. Software Models Simulation environment Target/constraints
3DOF dynamics model, rotating spheroidal earth model, 4th order gravitation model, MSISE 93 atmosphere model
Target/constraints
Velocity stopping condition
Altitude and flight path angle targets for scramjet burn only
Parameterised vertical acceleration profile

13. Common Mission Parameters
Booster 1 burn
10s burn, Alt =9.5km, Vel =550m/s
Coast
45.4s, Alt =15.9km, Vel =295m/s
Booster 2 burn
25s burn, Alt =19.6km, Vel =2411m/s
44.6s, Alt =25.3km, Vel =2000m/s
Orbital stage
Two burns, Alt =200km, Vel =7784m/s

14. Mission Profiles for Hydrogen Fuelled Vehicle

15. Hydrogen Case - Altitude Profile

16. Hydrogen Case - Velocity Profile

17. Mission Profiles for Hydrocarbon Fuelled Vehicle

18. Hydrocarbon Case - Altitude Profile

19. Hydrocarbon Case - Velocity Profile

20. Payload Estimation Mass and state at end of the scramjet burn
Scramjet mass fractions
Hydrogen fuelled waverider εpropellant = 0.58
Hydrocarbon fuelled waverider εpropellant = 0.7
Orbital stage
upper stage εstructure = 0.15
ΔV requirement based on Hohmann transfer

21. Mass Breakdown Hydrogen fuelled case Initial mass: 2000kg
Fuel mass: 316kg
Structure mass: 1000kg
Orbital stage mass: 684kg
Payload to 200km circular: 108.5kg
Payload mass fraction: 1.16%
Hydrocarbon fuelled case
Initial mass: 2000kg
Fuel mass: 258.8kg
Structure mass: 918.3kg
Orbital stage mass: 822.9kg
Payload to 200km circular: 36kg
Payload mass fraction: 0.38%

22. Discussion
Payload mass fractions similar to rockets even though much higher Isp?
Considerably lower fuel mass fractions
i.e. more of stage mass is structure, compared to rockets
Structure is more expensive than fuel.
These systems need to be reusable to be financially viable

23. Discussion Lighter scramjet stage for the hydrocarbon fuelled system
Hydrogen fuelled vehicle considerably higher payload capability than hydrocarbon fuelled case
Longer duration burn at higher Isp for H2 case
Better packing efficiency does not help the hydrocarbon vehicle as a large aerodynamic area is needed to maintain lift at high altitude so the vehicle cannot be made smaller.

24. Conclusions Similar payload mass fractions to rockets
Therefore need to be reusable
Hydrocarbon fuelled case has lighter structure than hydrogen fuelled case
Better packing efficiency
Better packing efficiency cannot be utilised due to aerodynamic requirements

25. Questions?