The Specific Impulse Potential of Hydrogen Scramjet
S.V. Horn, Senior AIAA Member, retired
Simplified model of the hydrogen scramjet was outlined which neglects a pre-combustion shock and
dissociation. Also, the inviscid inlet aerodynamics analysis is not complete, as well as the pressure drop
in combustor is not addressed in the simplified steady- state scramjet model. Thus simplified twodimensional (2-D) inviscid Excel model favors scramjet over rocket. The combustor model relied on the
moderate heat addition at equivalence ratio less than 0.6 as there is evidence of a complete combustion
based on reaction rates for hydrogen for a small scramjet. Finite rate chemistry for hydrogen scramjet of
reference work revealed nearly complete combustion, while for the hydrocarbon scramjet the heat
addition can be limited by reaction rates for smaller scramjets.
The inlet with a four oblique shock wave (OSW) design would result the inlet temperature exceeding the
ignition temperature of hydrogen, Tign = 845 K. The specific impulse of scramjet , based on the oxygen
and fuel flow rate, is lower than the specific impulse of a liquid propulsion rocket engine (LPRE) for
equivalent ratio φ = 0.4. Besides a complete combustion for such case, the combustor temperature is
kept at acceptable level.
Reference for the LPRE burning liquid hydrogen (LH2) and liquid oxygen (LOX) reveals specific impulse
Isp = 450 s [1]. The gas generator LPRE J-2 of Rocketdyne showed Isp = 435 s in vacuum, while the
staged combustion LPRE SSME showed Isp = 454 s in vacuum at 109% power level. The combustion
chamber temperature in SSME is Tc = 3600 K. Kerosene and LOX engine RD-170 of CIS had a tested
vacuum specific impulse Isp = 337 s [1]. Solid rocket motor (SRM) as ATK RSRM of Space Shuttle and
next generation Ares launch vehicle has specific impulse Isp ~ 278s in vacuum and the chamber
temperature Tc ~ 3600K.
Specific impulse in ramjets and scramjets historically relied on a different definition as flow rate of
oxidizer (air) is not included. The 2-D simplified model presented here provided comparison of specific
impulse normalized by oxygen and fuel flow rate. Such a comparison formed a common base, yet
advantage of scramjet vehicle, not carrying oxidizer, was fully recognized. A hydrogen scramjet
potential was estimated as based on the historic definition specific impulse of Heiser and Pratt [2, 3].
Current evaluation of historical specific impulse included fuel mass fraction [4].
Hypersonic vehicle with ramjet and scramjet propulsion would operate in a wide range of the unsteady
states. Heat addition limit was based on continuity equation between inlet and combustor. Excessive
heat addition would cause the subsonic combustion or transonic combustion as flow non-uniformity in
vicinity of injectors and propagated mixing and combustion regions is large. Simplified approach
neglected flow separation in isolator and upstream effect of combustor.
Combustor model of Roberts [4] provided plausible results, namely with the friction included in model.
The specific impulse was defined as the specific thrust normalized by air flow rate. Effects of a precombustion shock wave train (PCST) and dissociation were included in fifth reference [5]. The late
scramjet model, part of the presented model, featured no PCST in isolator. PCST pressure ratio was
shown dependent on the stagnation temperature ratio for case of the combustor Mach number M4 =1.
Dissociation effect was found to have a strong impact on the stagnation temperature reduction at the
free-stream Mach number M > 8 or the combustion temperature exceeding 4000 K. It included
dissociation of the inlet compression system evaluated from the single normal shock as worst case.
Further dissociation takes place in combustor, yet recombination in exhaust nozzle reduces overall
dissociation [5]. For the lower Mach number M < 8, the stagnation temperature reduction was found
lower than 5 %. Dissociation of oxygen and hydrogen starts at temperature 2500 K (M~8) and
dissociation of nitrogen starts at 4000 K (M~10). For inlet dissociation, shown on Fig.2, relevant is
dissociation of oxygen and nitrogen, dissociation in combustor will include hydrogen, as well as for
recombination in nozzle.
Reaction rate modeling revealed a near complete combustion up to equivalence ratio 0.6 for hydrogen
combustion. Hydrocarbon scramjet burns JP-7 which cracks to ethylene C2H4 and methane CH4. It was
estimated that the hydrocarbon combustion would not lead to a complete combustion in a given
scramjet, even at low equivalence ratio [5]. Continued development of scramjet model employed the
flamelet concept to avoid an intense use of CFD, namely during design studies [6]. Computational fluid
dynamic (CFD) tools as VULCAN and CFD++ were used extensively in scramjet development [7]. The
combined equivalence ratio from 2 injector locations approached φ = 1 both in tests and CFD runs.
History of scramjet development was shown by Curran [8]. Major challenges in scramjet development
are the high temperature at leading edge and combustor as well as the high dynamic pressure at high
hypersonic Mach number. Heiser and Pratt mentioned difficulties to design space vehicle for the
dynamic pressure in excess of 2000 psf (95 kPa) [3].
Example of the scramjet inlet which was designed with three to six oblique shock waves (OSW) is in
reference [4]. An inviscid flow deceleration for the inlet with three OSW was deceleration D = M3/Mo,
D = 0.37 and D = 0.31 for the six OSW inlet.
Heiser and Pratt [8] showed correlation of the inlet efficiency of Waltrup, decreasing with deceleration D
.
Good deceleration and high inlet efficiency was achieved in REST hydrocarbon engine [9] from Mo = 6 to
M2 = 3.74 (inviscid) or M2 = 3.28 (viscous analysis). Deceleration was D = 0.62 or 0.55. It is a sufficient
deceleration for JP-7, the ignition temperature for JP-7 is Tign = 514 K [4], yet not for hydrogen. REST
hydrocarbon engine is designed for Mach number Md =7.1 and is to be operated in range of free-stream
Mach number Mo = 4.8 to 8. Major impediment of the hydrocarbon scramjet is reaction rate which is
ten times slower than for the hydrogen scramjet [5, 11]. Despite the odds, X-51 demonstrated the
mission acceleration a = 0.18 g [12]. Compare it with the acceleration of the solid rocket motor (SRM)
rocket, it is usually limited to a = 6 g for structural reasons.
Hydrocarbon scramjet studies with the 2-D simplified model, which includes friction and injector velocity
[4], were based on assumption that combustion was complete, namely for long scramjet vehicle. The
scramjet design employed in the presented 2-D inviscid Excel model was 13.2 meters long [4]. Residence
time is higher than 5 ms. X-43A scramjet is 3.7 meters long, so is its residence time of air and fuel
particles is below 5 ms. Combustion can be completed in larger scramjet. Hydrocarbon scramjet vehicle
X-51 is 4.27 m long.
Table 1. Specific Impulse of Hydrogen Scramjet for the Four OSW Inlet
FER, Fuel equivalence ratio φ = 1
Mach
5.7
6
7
8
T3
[K]
857
893
1012
1106
T3/To
Tt3
Tt4
Tt4/Tt3
3.78
3.93
4.46
4.87
1538
1680
2225
2838
3367
3498
4000
4561
2.19
2.08
1.80
1.61
Isp
(O2+fuel)
333
316
262
215
Isp
[3,4]
2616
2460
1985
1594
Isp
(fuel)
2714
2574
2136
1745
o – free-stream, t - stagnation state, 3 - combustor entry, 4 - combustor exit
Equivalence Ratio φ = 1 may not guarantee complete combustion
Isp(O2+fuel) = thrust / oxygen and fuel weight flow rate
Isp(fuel) = thrust / fuel weight flow rate
Note that above Mach 8, dissociation reduces Tt more than 5% [5], see Fig.2.
Table 2. Mach numbers and inlet temperature for the four OSW inlet
Mach
M isolator
Mcombustor
Mnozzle
5.7
2.1
1.13
4.22
6
2.21
1.21
4.24
7
2.57
1.47
4.36
8
2.94
1.72
4.53
The maximum heat addition, or the stagnation temperature ratio of combustor, to keep the supersonic
combustion is on the Fig.1 for the specific heat ratio k = 1.24 and 1.36. It is based on the continuity
equation between the inlet at cowl and the combustor exit.
Effect of dissociation on the reduction of the stagnation temperature is on Fig. 2. The minimum
temperature at inlet to achieve hydrogen (excluding silane SiH4) ignition is Tign = 845 K. To achieve a
fast ignition of SiH4 (less than 50 microseconds at 1.25 atm) the inlet temperature should be Tinlet ~
1000 K [10].
The latest example is the scramjet vehicle designed for Mach number Mo = 8 [18]. The design
equivalence ratio FER = 0.5. Performance data predicted by MASIV are available both for FER = 0.5 and
0.3. Specific impulse Isp(fuel) = 1866 s and 2183 s respectively. Specific impulse Isp(O2+fuel) = 133 s and
96 s for FER =0.5 and 0.3 respectively. Scramjet is 24.4 m long and estimated residence time is 10.3 ms.
Deceleration of inlet was estimated at D = M3/Mo ~0.5, from Mo = 8 to M3 ~ 4, insufficient for the
hydrogen ignition without additives. Estimated power of scramjet is about 1 GW at Mach number Mo=8.
The simplified 2-D steady state Excel model used scramjet designed for Mach number Mo=4 [4]. The
specific impulse at Mach number Mo=8 was evaluated at Isp(O2+fuel) = 84 s and Isp(fuel)=1183s.
Summary
It is difficult to use air as oxidizer as it is done in ramjets and scramjets. Benefit of not carrying oxidizer
in space vehicle or missile seems to be great, yet in scramjet the combustion temperature can be
higher than in rockets, LPRE or SRM, by 1000 K. Pure oxygen as oxidizer (LPRE Isp ~450s) is better than
air or oxidizer in SRM as ammonium perchlorate NH4ClO3, where specific impulse Isp ~ 300 s. Specific
impulse of Heiser and Pratt [3] updated by fuel fraction [4] was in qualitative agreement with specific
impulse
Isp(fuel) = Thrust/weight flow rate of fuel.
It is expected that TBCC (turbine based combined cycle) and scramjet would provide propulsion for
space vehicle in 2030, in 70 years since scramjet development [13]. One of challenges is to provide
uniform inlet flow field to TBCC.
For a space access reduction of oxidizer mass in scramjet is partially lost with mass of the scramjet
itself. TBCC with scramjet is concept inferior to rocket space access today.
Scramjet may not be suitable as missile as well as short missile does not provide a sufficient residence
time for complete combustion. Yet for tomorrow it is imperative to match or exceed capability of
hypersonic missile BrahMos and establish air and naval dominance. Indian and Russian sea skimming
cruise missile BrahMos is expected to achieve Mach number M ~ 5 to 7 in year of 2017, current Mach
number with rocket and ramjet propulsion is M ~ 2.8. For better atomization of LH2, kerosene is
beneficial (Kumar, DRDL), also reported by University of Munich, Germany.
Specific impulse based on thrust normalized by the oxygen and fuel weight flow rate was estimated for
hydrogen scramjet. Inlet design with the four OSW would secure the ignition temperature for hydrogen
845 K. It would also assure low ignition times for silane, less than 50 microseconds. The specific impulse
of scramjet is lower than the specific impulse of rocket. Fuel equivalent ratio (FER) φ =1 resulted in
results shown in Table 1. Also the inlet conditions correspond to altitude at 25 to 31 km, to keep the
dynamic pressure of the free-stream air at Pd = 1000 psf [4]. To reduce the combustor temperature to
Tt4=Tc < 4000K, FER = 0.4 is needed. Isp (O2+fuel)~100 at Mach number Mo = 6 and FER = 0.4. The
latest scramjet design by MASIV [19] would achieve Isp(O2+fuel)=133 s at Mach number Mo=8 and FER
=0.5. The specific heat ratio, molecular weight, and gas constant were updated for given FER by CEA
code of the NASA GRC [ 14] and the altitude specific heat ratio k=1.36 and the lowest specific heat ratio
k = 1.24 are reflected in Fig.1.
Dissociation of oxygen at 2500 K (M~8) and nitrogen at temperature 4000 K (M~10), would reduce the
stagnation temperature of combustor more than 5 % (see Fig.2) and specific thrust at the free-stream
Mach number Mo = 8. Further dissociation in combustor and its reduction in exhaust nozzle by
recombination would provide overall dissociation [5].
Scramjet inlet can be unstarted as in supersonic flow any perturbation can propagate upstream.
Perturbations can be heat addition inducing shock waves in combustor or the injection of fuel. Fuel
injection effect studied in experimental test rig by Planar Laser Rayleigh Scattering detected pseudoshock leading to a complete unstart potential [15]. Rocket, SRM or LPRE, cannot exhibit such
phenomenon caused by open inlet of scramjet. Unstart probability is being addressed Stanford
University team at PSAAP.stanford.edu website. Supersonic propagation from external surfaces of the
hypersonic vehicle is coupled with the internal aerodynamics of scramjet which would impact thrust
level as well. It is possible that an external aerodynamics wave can unstart scramjet inlet as well.
Heat transfer in scramjet is one of key challenges as it impacts heat addition and aerodynamics. High
temperature materials as Inconel and Haynes alloys and carbon composites are to be employed with
advanced cooling concepts. Example of software for ramjet and scramjet heat transfer is SRHEAT™ ,
available by Spiritech [16]. Multidiscipline Design Optimization (MDO) is being used by Boeing for
development of scramjet hypersonic vehicles, like X-43A, currently flight tested by launch from aircraft
B-52B and propelled by ATK SRM Orion50S to hypersonic speed [17]. Current materials are recognized
to be suitable for temperatures up to 3000 K (5000°F) [17].
Scramjets were already in development for 50 years. Major breakthrough can reduce cost of launch to
altitude above an Earth thermosphere when using scramjet. Wide spread of scramjet as weapon would
contribute to minor oxygen depletion on Earth, scramjet power is of order 1 GW for compared scramjets
24.4 m and 13.2 m long, at Mach number Mo=8. Speed of dissociation reactions was also indicated at
[19], besides studies in [5]. The most recently, the scramjet inlets with 3 OSW and 4 OSW of Roberts [4]
were analyzed with CFD Fluent code [20]. Viscous deceleration D was close to inviscid one, yet the
pressure and temperature ratii were smaller than the inviscid ones. Mass averaging of temperature and
pressure would improve it.
References
[1] G. P. Sutton: Rocket Propulsion Elements, Wiley&Sons, 1992, Sixth Edition.
[2] R.S. Fry: A Century of Ramjet Technology Evolution, Journal of Propulsion and Power, 2004.
[3] W.H. Heiser, D.T. Pratt: Hypersonic Airbreathing Propulsion, AIAA 1994.
[4] K.N. Roberts, D. R. Wilson: Analysis and Design of Hypersonic Scramjet Engine with a Transition Mach
Number 4.00, AIAA-2009-1255. Includes MSc. Thesis of K.N. Roberts.
[5] S.M. Torrez, N.A. Scholten, D.M. Micka, J.F. Driscoll, M.A. Bolender, D.B. Doman, M.W.
Oppenheimer: A Scramjet Engine Model Including Effects of Pre-combustion Shocks and Dissociation,
AIAA-2008-4619.
[6] S.M. Torrez, J.F. Driscoll, D.D. Dalle, D.J. Micka: Scramjet Engine Model MASIV: Role of Mixing,
Chemistry and Wave Interactions, AIAA-2009-4939.
[7] K. Cabell, N. Haas, A. Storch, M. Gruber: HIFiRE Direct-Connect Rig (HDCR) Phase I Scramjet Test
Results from the NASA Langley Arc-Jet Heated Test Facility.
[8] E. Curran: Scramjet Engines: First Forty Years, JPP, Vol.17, No.6, 2001.
[9] M.K. Smart: Design of 3-D Hypersonic Inlets with Rectangular-to Elliptic Shape Transition, JPP 1999,
Vol.15, No.3.
[10] M. Koshi et al. : Chemical Kinetics of Silane Combustion, J. of Phys. Chem. 1997.
[11] W. Huang et al.: Hydrogen Fueled Scramjet Combustor – the Impact of Fuel Injection,
www.intechopen.com
[12] R. Mutzman, S. Murphy, X-51 Development: A Chief Engineer’s Perspective, AFRL 2011, 17th AIAA
International Space Planes and Hypersonics Systems and Technologies Conference.
[13] B. Thieman: Next Generation Space Access, Oct.19, 2010, AFRL.
[14] S. Gordon, B.J. McBride: Computer Program for Calculation of Complex Chemical Equilibrium
Compositions and Applications, NASA-TM-1311, 1994.
[15] H.Do, S-K. Im, M.G. Mungal, M. A. Capelli: Supersonic Duct Unstart Induced by Fuel Injection.
Stanford University Report TSFP7, 2011, submitted to Experiments in Fluids.
[16] E.J. Gamble, J. Gutierrez, J. Bachmann, T. Jobin, D. Williford, C.N. Raffoul: Development of
Scramjet/Ramjet Heat Transfer Analysis Tool (SRHEAT™), AIAA-2008-4614.
[17] K. Bowcutt: Advancing Promise of Hypersonics Through Flight Testing, AIAA Orange County Section
Speaker Series, 2013.
[18] S.M. Torrez, D.D. Dalle, J.F.Driscoll: Dual Mode Scramjet Design to Achieve Improved Operational
Stability, AIAA-2010-6957.
[19] R.A.Strehlow: Combustion Fundamentals, McGraw Hill, 1984.
[20] A. Siddiqui, G.M. Ahmed: Design and Analysis of Scramjet Engine Inlet, International Journal of
Scientific and Research Publications, January 2013.
Fig.1 Heat Addition Limit to keep Mach number in Combustor M = 1
Thermal Choking and Unstart Avoidance
Continuity Condition between the inlet cowl and the combustor exit
t - stagnation state, 3 - combustor entry, 4 - combustor exit
Heat Addition Limit of Supersonic Combustor
k = 1.24 and 1.36
3.0000
2.5000
2.0000
Tt4/Tt3 1.5000
1.0000
0.5000
0.0000
1
2
3
4
5
6
7
Inlet Mach Number at Cowl
8
9
10
Fig.2 Reduction of the Stagnation Temperature by Dissociation in Inlet [5]
Presented with permission
Reduction of the stagnation temperature by
dissociation in inlet compression system
0.3
0.25
0.2
ΔTt/Tt 0.15
0.1
0.05
0
4
6
8
10
12
Free-stream Mach number
14
16