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

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Presentation on theme: "Orbital Payload Delivery Using Hydrogen and Hydrocarbon Fuelled Scramjet Engines M. R. Tetlow and C.J. Doolan School on Mechanical Engineering The University."— Presentation transcript:

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 –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 Ignition Delay Heat Transfer Shear Stress Displacement Thickness Growth Combustion Area Change Injector Flow From Inlet

9 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. Quasi-1D Scramjet Propulsion Model

10 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). T4 Experiment Parallel Combustor 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 –3DOF dynamics model, rotating spheroidal earth model, 4 th 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  Coast –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 ε propellant = 0.58 –Hydrocarbon fuelled waverider ε propellant = 0.7  Orbital stage –upper stage ε structure = 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?


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