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With advances in chemical engineering, could chemical propulsion have a fuel (or fuels) that would allow it to compete with newer, more advanced form of propulsion (such as electrical or nuclear) or are the chemicals being used now likely to be as good as it's going to get?

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Exceeding the specific impulse (i.e. mass efficiency) of electrical or nuclear propulsion is unlikely with chemical propulsion.

Practical chemical rockets top out around 460 seconds Isp, with exotic, but impractical, propellants going up into the 500s.

Nuclear-thermal rockets (using the heat of a controlled fission reaction to accelerate hydrogen or other propellants) were expected to deliver better than 800 seconds Isp. Such engines were developed (NERVA, RD-0410) but never flown.

Electrical thrusters (Hall effect, ion, VASIMR, etc) produce Isp in the 2000-5000 second range.

Competing on thrust instead of specific impulse, of course, the rankings are reversed; electrical thrusters have extremely low thrust-to-mass ratios, and chemical propellants very high ratios, with NTR in between.

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  • $\begingroup$ "Competing on thrust instead of specific impulse, of course, the rankings are reversed; electrical thrusters have extremely low thrust-to-mass ratios, and chemical propellants very high ratios, with NTR in between." Not necessarily (although it is true for essentially all current designs); some proposed designs (such as nuclear saltwater rockets and some gas-core NTRs) manage to combine both high thrust and high specific impulse. $\endgroup$ – Sean Mar 26 '18 at 1:05
  • $\begingroup$ Has enough thermal analysis been done to suggest that NSWRs are actually possible without melting the nozzle? And has anybody actually proposed a gas-core NTR that counts as high thrust rather than moderate thrust? $\endgroup$ – ikrase 10 hours ago
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Honestly, the efficiency of chemical rockets we use today are nearly maxed out for practical fuel combinations. Also, chemical rocket engines will most likely never reach the efficiency metrics of electrical and nuclear propulsion systems (the propulsive energy conversion is fundamentally different). However, when you need a substantial amount of thrust, the thrust to weight ratio of chemical propulsion systems dominates the playing field.

The combustion process used today in all rockets/gas-turbines is called deflagration, which is a relatively slow release of heat and a constant pressure combustion process. The only thing that can be done to raise the $I_{sp}$ for chemical propulsion systems (and is currently being studied) is detonating the fuel, which is a pressure gain combustion process. Detonation based propulsion systems come in two varieties (at least at the research and experimental level). The first is called the pulsed detonation engine (PDE), which has received a great deal of treatment since the early 90s. The second is the rotating detonation engine (RDE), which is heavily active in current aerospace research.

Here is a plot from an analytical study (Morris, C. I., “Numerical Modeling of Single-Pulse Gasdynamics and Performance of Pulse Detonation Rocket Engines,” Journal of Propulsion and Power, Vol. 21, No. 3, 2005, pp. 527-538.) that details the idealized performance of a single-cycle pulse detonation rocket engine vs the conventional steady state rocket engine (SSRE) for oxyhydrogen fuel. Note, the blowdown pressure ratio does not go to infinity (i.e. a vacuum) so comparison with the vacuum level performance is not reported, however, the idealized PDRE is more efficient than the conventional rocket engine over the entire operational blowdown pressure ratio range. Now granted this is all idealized performance, and harnessing the energy of detonation wave for thrust applications is generally very challenging. Here are some attempts in the laboratory setting at the academic level. This testing at the academic level is conducted by PhD students, however more sophisticated operations exist that are not open source at facilities like the Air Force Research Laboratory.

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The RDE is a more sophisticated concept for a detonation based engine. It is more practical in application than the PDE, due to its non-pulse operation. Here is a schematic of the operation of the RDE.

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The RDE is actively being researched at many universities and industry companies (Aerojet/Rocketdyne and the Aerospace corporation to name a few). The primary goal as of now (economically convincing for funding) is to integrate the PDE/RDE into next generation gas turbine engines. However, the rocket application remains wide open given the maturing technology. Here are some videos of RDE testing at various universities:

1.) University of Texas at Arlington
2.) Purdue University
3.) University of Cincinnati

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  • $\begingroup$ That chart looks to me like the high end of detonation engines would be maybe another 20 seconds of Isp, i.e. topping out around ~480s for a very high expansion ratio hydrogen engine, which wouldn't compete with NTR or electrical propulsion. $\endgroup$ – Russell Borogove Jan 9 '17 at 2:30
  • $\begingroup$ I never claimed that a detonation based engine would compete with electric or nuclear propulsion systems regarding $I_{sp}$. The OP asked if there are chemical "improvements" that would better the operation of chemical rocket engines. Regarding chemical combustion, one can either use a more energetic fuel, or change the mode of combustion. I proposed the case of changing the mode of combustion from deflagration to detonation, which is actively being pursued as academic research PhD topics, and within the Air Force Research Laboratory. However, partially filled PDE combustors is an interesting $\endgroup$ – TRF Jan 9 '17 at 2:36
  • $\begingroup$ @RussellBorogove cont. concept where $I_{sp}$ can be raised to 3-5 times that of the baseline fully filled PDE ($I_{sp} \sim 600 - 1,000 $ sec). Here is a numerical study (arc.aiaa.org/doi/abs/10.2514/1.9514?journalCode=jpp), however, there are also many experimental studies with similar results. The partial fuel filling method is very useful for raising the $I_{sp}$ of a PDE, but you will suffer a loss in the overall attainable thrust. $\endgroup$ – TRF Jan 9 '17 at 2:41

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