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One important part of the design of a high performance powertrain is the choice of the propellant

  • type
    • monopropellant
      • fuel and
      • blend,
    • bipropellant
      • oxidizer and
      • monopropellant,
    • tripropellant,
  • transformation by
    • preprocessing
      • deeply cooling,
      • liquefaction,
      • heating,
      • decomposition and
      • reformation,
    • combustion,
  • integration into a system for
    • collection, separation and enrichment,
    • regenerative cooling,
    • thermal protection and
    • ionization.

    Thoughts, requirements and constraints
    To decide which propellant or combination of propellants fits best for an engine several thoughts, requirements and constraints have to be considered and, if needed, weighted against each other for realizing the highest performance. Some of the most interesting points we have found in the past while developing our Active Hypersonic En- gine are:

  • Precious O (oxygen) is wanted.
  • H (hydrogen) is easy to get, so that it's sometimes simply dumped overboard by some Air Collection and Enrichment Systems (ACES)/Li- quid Air Cycle Engine (LACE) systems.
  • H (hydrogen) is as propellant significantly more favored than unre- formed CH (hydrocarbon) based fuel, because it has a higher specific impulse (prefered at higher altitudes) and is an essential part of the cleanest combustions, while CH propellant has a higher thrust-to- mass (important for takeoff).

  • LOX/LH2 offers the highest Isp with 450 s of any realistic chemical propellant option. It also offers a relatively high oxidizer/fuel mass ratio of 1:6.
  • O2/CH4 offers an Isp of 385 s and a mixture ratio of 1:3.5.
  • O2/C3H4 is expected to reach an Isp of 370 s.
  • O2/RP-1 offers a specific impulse of 355 s with high specific impulses at mixture ratios of about 1:2.6.
  • H2O2/JP-5 offers an Isp of about 330 s and burns with a mixture ratio of 1:7.3.
  • H2O2/C2H6 (ethane) offers an Isp of around 307 s.
  • H2O2/C2H6-C3H8 mix offers an Isp of around 305 s.
  • H2O2/C3H8 (propane) offers an Isp of around 303 s.
  • H2O2/RP-1 offers an Isp between 299 - 303 s.

  • Some combinations of H2O2 and fuel equal or exceed the perform- ance of LOX/LH2.
  • Many propellant combinations beat LOX/LH2.

  • B (boron)/CH2 (approximate formula of kerosene) slurry with a mixture ratio of 55%:45% exhibits an Isp increase of 50 to 100% over kerosene in a ramjet, which then is close to the Isp of hydrogen in a ramjet.
  • O2/B/RP-1 is expected to have an Isp of 530 - 710 s in a rocket engine, but can make problems due to agglutination in the nozzle.

  • LH (liquid hydrogen) provides more than three times as much ener- gy and absorbs six times more heat per measure unit than any other fuel.
  • LCH4 (liquid methane) provides more energy and can absorb five times as much heat as CH2 (aprx. kerosene).

  • LH (liquid hydrogen) has a low density, which means larger fuel tanks, a larger airframe and more drag.
  • LCH4 (liquid methane) is three times denser than LH (liquid hydrogen).
  • With pre-chilled propellants at 10 K above their melting points, the increase in payload attributable to the higher density ranges from approximately 20 - 30%.
  • We don't want to use cryogenic states of propellants. Such a pro- pellant system has too much complexity and for example in the case of the extremely aggressive LH becomes very expensive. Also, a sy- stem with noncryogenic states of propellants can remain operational in orbit for a very long time without additional complexity.
  • CH2 (aprx. kerosene), CH3OH (methanol or methyl alcohol), H2O2 (hydrogen peroxide) and H2O (water) are noncryogenic (liquid at room temperature).

  • LCH4 (liquid methane) is easier to handle compared with LH (liquid hydrogen).
  • LCH4 (liquid methane) is widely available compared with CH2 (aprx. kerosene).

  • LH (liquid hydrogen) has a heat of combustion of 51,800 Btu/lbm.
  • LCH4 (liquid methane) has a heat of combustion of 21,500 Btu/lbm.
  • CH2 (aprx. kerosene) based JP has a heat of combustion of around 19,300 Btu/lbm.
  • C7H14 (methylcyclohexane) has a heat of combustion of 18,800 - 19,500 Btu/lbm.
  • CH2 (aprx. kerosene) reformation into high hydrogen fraction fuel with a heat of combustion between 20,200 Btu/lbm and 30,000 Btu/lbm for Russian CH2.
  • CH4 (methane) reformation into high hydrogen fraction fuel with a heat of combustion of 59,600 Btu/lbm.
  • The heat capacity of some of the reformed CH (hydrocarbon) pro- pellants can be higher than H (hydrogen).
  • CH (hydrocarbon) propellant is cheap. But the reformation of CH based fuel needs precious O and gives CO or CO2. On the other hand we have an air-breathing engine and maybe still not enough O.

  • If LO2 (liquid oxygen) is needed for combustion, then it's the majority of the weight of a spacecraft on liftoff. So if some of this can be collected from the air on the way, it might dramatically lower the take-off weight of the spacecraft.

  • Decomposed H2O2 used as an oxidizer gives a somewhat lower specific impulse (as high as 350 s with CH2 (aprx. kerosene) based RP-1) than LO2 (liquid oxygen), but is dense, storable, noncryogenic and can be more easily used to drive gas turbines to give high pressures using an efficient closed cycle.

  • H2O2 in 98% concentration is 1.43 times as dense as water.
  • H2O2 has superior heat characteristics and may also be used for regenerative cooling.
  • We have enough heat and pressure "for free" very quickly and most of the time while flying in the atmosphere of a planet, so that H2O, H2O2 and every mixture of both can be decomposed, and CH fuel be reformed.
  • If the H2O2 and H2O give the oxidizer O, and the H2 is the fuel, then we have here H2O (water) as fuel and the most clean combu- stion with H2O and a little bit O2 at the end.

  • Further propellants, blends and combinations like
    • CH3OH (CH4O or MeOH; methanol or methyl alcohol),
    • CH3CH2OH (C2H5OH, C2H6O or EtOH; ethanol or ethyl alcohol),
      • E85 (85% ethanol and 15% gasoline)
    • C2H6 (ethane),
    • CH3OCH3 (C2H6O; dimethyl ether),
    • HOCH2CH2OH (C2H6O2; ethylene glycol),
    • C3H4 (methylacetylene),
    • C3H4O (propargyl alcohol),
    • C3H6 (propylene or propene),
    • C3H8 (propane),
    • CH3CHCH2O (C3H6O; propylene oxide),
    • CH3CH2CHO (C3H6O; propionaldehyde or propanal),
    • CH2CHCH2OH (C3H6O; ally alcohol),
    • CH3CHOHCH2OH (C3H8O2; propylene glycol),
    • CH3CH2OCH2CH3 (C4H10O, (C2H5)2O or Et2O; diethyl ether or ether),
    • C#H#O#,
    • C7H14 (MCH; methylcyclohexane; endothermal),
    • C#H#, as well as
    • other evironmental friendly propellants
    are appliable.

    Rocket-Based Combined Cycle Propellant
    Rocket-Based Combined Cycles can be powered by using hydrogen peroxide as the oxidizer of bipropellants.

    Hybrid rocket engines can be powered with a bipropellant like

  • H2O2.H2O/HTPB (hydroxyl-terminated polybutadiene), for example.

    Liquid rocket engines can be powered with the same propellants as scramjet engines like

  • hydrogen oxide based propellants
    • H2O (water),
  • hydrocarbon based propellants
    • CH2 (aprx. kerosene) based
      • Rocket Propellant (RP-1) and
      • Jet Propellant (Jet-A and Jet-A1, and JP-5, JP-7 and JP-8),
    • CH4 (methane),
    • C2H6 (ethane),
    • C3H8 (propane),
    • C7H14 (MCH; methylcyclohexane),
  • hydrocarbon based propellant blends,
  • hydroxyl groups based propellants
    • CH3OH (CH4O or MeOH; methanol or methyl alcohol),
    • CH3CH2OH (C2H5OH or EtOH; ethanol or ethyl alcohol) based
      • E85 (85% ethanol and 15% gasoline),
    • HOCH2CH2OH (C2H6O2; ethylene glycol),
    • C3H4O (propargyl alcohol),
    • CH2CHCH2OH (C3H6O; ally alcohol),
    • CH3CHOHCH2OH (C3H8O2; propylene glycol),
    or alternatively
  • ether based propellants
    • CH3OCH3 (C2H6O; dimethyl ether),
    • CH3CH2OCH2CH3 (C4H10O, (C2H5)2O or Et2O; diethyl ether or ether).

    Exemplary Choices and Integration

  • Bristol Siddeley Gamma used H2O2 and CH2, which is the second most clean combustion with H2O (water) and a little bit CO2 (carbon dioxid) at the end.
  • Leninets has proposed to use H2O and C#H# with reformation, which as air-breathing engine and with water-gas shift reaction for CO reduction would be as well the second most clean combustion with water and a little bit CO2 at the end.
  • We do want the closed cycle, but like the expander cycle (a linear aerospike engine does not suffer form expander cycle limitation due to the fact that the linear shape of such an engine is not subject to the square-cube law), and a little bit of the gas generator cycle and staged combustion cycle without expensive turbopumps. This leads to the middle of all, that has much in common with the integration of the concepts by Bristol Siddeley Gamma and Leninets: H2O2 and O of decomposed H2O as oxidizer, and H of decomposed H2O as well as reformed C#H# and C#H#O# instead of unreformed CH2 (aprx. kerosene) as fuel. H2O is already there. Both the oxidizer and the fuel are used for regenerative cooling as part of a thermal protection system.

    Exemplary Application
    Ideally, a propulsion system is powered by one propellant only to keep complexity of the fuel system as low as possible. But that is not generally a precondition for an engine concept.

    The rule of thumb for the development of our engines is to take the optimal steps for the boot(strap) process of an engine and its integration into one powertrain by finding the right balance between H2O2, H2O and the propellant, and by simultaneously reducing the complexity, especially as much as possible at the sections where rotating hardware is used to pump the propellant for the whole operational range.

    One initial step can be to burn a hydrocarbon propellant to produce heat and pressure for the extraction and decomposition of the H2O2 (hydrogen peroxide) while the engine and the thermal management system of the vehicle are cold.

    Then all the heat needed for driving the turbopumps and all of the propellant for the first/start phase (ramrocket and rocket ejector mode) exist.

    In the later phase while flying above a velocity of 1.5 Mach enough heat, and hereby H2O steam, and electricity generated by thermal and/or kinetic energy recovery systems is available for the propell- ant preprocessings.
    Furthermore, due to the high temperatures the integration of the propellants into additional systems to take more advantages is only a logical step. A very interesting example is the ionization of the propellant before it enters the engine and the acceleration of the still ionized exhaust at the nozzle of the engine by applying the technologies of ion thrusters.

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    Christian Stroetmann GmbH