Reusability of the rocket engines, which are the most expensive part of the rocket, is very important to keep low cost of the space rockets and in this way making reality the space programs considered expensive. But does this reusability affect the performance of the rocket engine?

I am interested to know the performances of the same type of rocket engines but in different versions, one reusable and the other single launch. Probably reusability affects the weight (in this way even thrust to weight ratio) and the cost of the engine but what about other elements that have to do with performance such as Thrust it's self, Specific Impulse, Chamber Pressure. If performances are different how much could be different. It is about something significant let say a considerable different performance, or a minor influence in performance almost identical. Illustrating with values or examples would be better to understand how much are the differences in case that differences exist.

EDIT: As Russell has mentioned a reusable rocket engine has to be more robustly engineered in general.

One of the problems at rocket engines is heating up from burning the propellants. By forcing more the engine the faster it's heating up and in a higher level. For a reusable engine if the temperature remains at those levels that are not dangerous, I guess it could operate for a longer time than single use engine. In the safe temperature level where the engine remains reusable, it could operate “indefinitely” (for a very long time). But I was wondering if the performance could be higher by forcing more this reusable engine, forcing it in those temperature levels that doesn’t destroy the entire engine but destroy essential parts and making its durability in that situation that can't survive two launches, as in the single use engines, or it is not so simple that you can force the engine as you wish and rise the Thrust, Specific Impulse or Chamber Pressure as you want.

So does it mean this that reusable engine has more " space " to force it at its limits in maximum, maybe with the cost that can’t survive another launch such as the single use engine design, but with the “profit” that the performance will be higher?

  • $\begingroup$ For RTLS (Return To Launch Site), a booster needs to kill eastward velocity and do a westward burn. This has made for a more vertical ascent profile. I expect this impacts payload mass more than engine design. $\endgroup$
    – HopDavid
    Commented Apr 28, 2016 at 13:16
  • 2
    $\begingroup$ If anything, reusability will positively affect the performance: solutions too expensive to be expendable will be employed. SSME with its four turbopumps, two preburners and a slew of other extremely advanced parts, resulting in performance about as close as possible to optimal for this fuel - being a good example $\endgroup$
    – SF.
    Commented Jul 19, 2016 at 17:30
  • $\begingroup$ Presumably we should be restricted to engines already "restartable" for a fair comparison? $\endgroup$
    – uhoh
    Commented Jul 24, 2016 at 4:17
  • $\begingroup$ @SF. You are right that if the objectives are harder such as reusable and advanced engine,than even the product(engine in this case)will be with high performance or bringing something new proving that could be done. Of course the SSME was an example(is not about only SSME),but who knows if the SSME would have a higher performance in a single use version(for this i am interested to know).So with the same"philosophy",science and the way that will operate,but without being restricted by the objective that should be reusable unconditional.For SSME was primary the reusability than the performance. $\endgroup$
    – Mark777
    Commented Jul 24, 2016 at 4:56
  • $\begingroup$ @uhoh it could be as a good reference, if you find the right comparison would be ok but i haven't find yet an engine model with the same design (of course those parts that should be reusable will be different, but the design at all and the way of operating to be the same) but one reusable and one single use and where in both cases the highest posible performance to be the primary objective. For example the RS-25E or F are single use but not a good example because their primary objective is to be as cheap as posible. If you could explain(answer) in theory or "engineering" way is even better. $\endgroup$
    – Mark777
    Commented Jul 24, 2016 at 5:15

5 Answers 5


A reusable engine has to be more robustly engineered in general, but as evidenced by SpaceX's Merlin engine, high thrust-to-weight ratios are certainly achievable in reusable engines.

Some engines, like the RS-68, use an ablative coating in the nozzle to dissipate heat, rather than active cooling in the nozzle walls. In the RS-68, this is actually heavier than the alternative -- thus yielding lower thrust-to-weight ratio -- but much cheaper to build.

I think a sufficient engineering and manufacturing budget can overcome any given performance limitation of a reusable engine design.

  • $\begingroup$ What about SSME if they would be single launch would it give more room engineers to push more performance specifications so to have higher thrust, higher specific impulse, higher chamber pressure, without changing it's concept, design for the components combination, and operating cycle with it's acoustic resonance chambers. I am not speaking for RS-25E or RS-25F which are single launch, but this versions are focused more to be low cost or simpler and cheaper, than to be a rocket engine single launch following SSME philosophy and with the highest possible performance no matter what the cost. $\endgroup$
    – Mark777
    Commented Apr 27, 2016 at 21:57
  • $\begingroup$ Doubtful. The SSME required extensive overhaul between uses. The Energia launcher's RD-0120, with no requirement for reusability, had barely better specific impulse and substantially lower thrust. $\endgroup$ Commented Apr 27, 2016 at 22:29
  • $\begingroup$ I see that you are not sure for higher performance part. But for RD-0120 part, it has not to do with SSME. They are both LOX/LH2, staged combustion chamber and have some similarities but they are different in design, concept, components, different way to make a functional staged combustion chamber. They have different choices for turbo-pupms and shaft for each fuel component, different channel-wall nozzle and cooling system, so not to make a comparison with RD-0120 single launch. A comparison of the SSME reusable with SSME single launch so something that has the same design. $\endgroup$
    – Mark777
    Commented Apr 27, 2016 at 23:24
  • $\begingroup$ The engineering choices in rocket engine design are extremely complex and often interrelated. It's never going to be as simple as taking expendable rocket A and "adding the reusabilty feature" or taking rocket B and removing the reusabilty. Merlin and SSME demonstrate that extremely good reusable engines are possible; everything else depends on how much budget you have. I don't believe reusabilty requires any inherent performance compromises. $\endgroup$ Commented Apr 28, 2016 at 2:49
  • 1
    $\begingroup$ There's also the matter of "degree of reusability." Some rocket engines would be perfectly reusable if they landed instead of crashing. Some reusable engines require costly inspections and frequent replacements. Some engines are strictly single-use, essential parts destroyed when the engine is fired, or the durabilities calculated such that the engine can't survive two launches. $\endgroup$
    – SF.
    Commented Apr 28, 2016 at 12:03

If you were to design 2 engines (one reusable, one expendable) with the same Thrust, Specific Impulse and Chamber Pressure:

the reusable engine would be heavier, more expensive, or both. For a reusable engine, parts have to be designed with a much longer lifetime than for an expendable engine. You can do that in two ways: make the part heavier to provide more wear margin, or use better materials that wear out more slowly.

The reusable engine would probably be more complex too: you'd want to design it to be easily accessible for inspection and (if necessary) part replacement. So you'd have access panels, piping with bolted flanges instead of welds etc. Maybe you'd move the turbopump to a place that's easily accessible (which needs more piping than the usual place on top of the combustion chamber).

When you want to increase performance of an existing engine, you have only a few options. You can make combustion more efficient (unlikely, because you already have efficient combustion to start with). Or you can inject more fuel and oxidiser. When you do that, combustion chamber pressure increases. Pressure in the turbopumps and piping increases as well.

These components are built to withstand a certain pressure. This design limit is a bit higher than the normal operating pressure. This is the margin you can work with.
Materials have a certain tensile strength, this depends on their temperature. In a rocket engine, components have to be kept below the temperature at which a metal would start to soften. Above this temperature, the component will break down rapidly.

If you take a reusable engine, and use it in expendable mode, the only way you can improve performance is by using this margin. If the engine has to run for 2 minutes, you can try and run the engine so that the components reach their critical temperature at 2 minutes 15 seconds, when the stage has burned out and the second stage is far enough away that a first stage explosion won't damage the second stage.

This is a difficult proposition though. When a material approaches its limits, its behavior is not linear. Say you've got an engine that can run indefinitely at 1000 ºC. If you raise its operating temperature by 5%, engine life will not drop by 5%. It might still be good for indefinite operation at the new temperature, or it might melt after 30 seconds. This depends on the detailed design of the engine and the margins that were used in its design.

The critical parts aren't wear parts: a combustion chamber is not designed to lose material as it is used (ablative lining, for example). If this were the case, it'd be easy: just run the engine to burn off the ablative lining in one go instead of making it last 10 missions.

How much performance you can gain this way depends on the exact engine design: the materials that were used, the margins used by the designers. My guess is, in a mature engine design you'll have very little margin to work with. Too much margin means the engine is too heavy.

Here's a paper that gives some insight into the design decisions that were made to make the SSME reusable. Reading the paper, I get the impression that the main differences were in the design process itself: they had to do research to find materials and component detail design that would be durable. So the design process was more expensive because they'd need to advance the state of the art, finding more reliable materials and construction methods.

They also needed to design the engine for maintainability. They needed to make sure wear parts were easily accessible to minimize the time spent on maintenance, they had to provide inspection panels, create tooling etc.

The Orbiter main engine is the first large liquid rocket engine designed specifically for a long service life. The engine is capable of completing 100 starts or 7.5 hours of operation between over­hauls. ... Specific changes to earlier engine designs were necessary to fulfill this long service life requirement. However, in each case, the weight of the design change was evaluated for its impact on engine weight-to-orbiter payload capability. Some of the engine systems and components designed for long service life are the hot-gas system, turbomachinery, and valve seats.

All turbopump seals operate with a positive clearance to prevent wear and ensure long life. Low bearing loads are ensured by a balance piston system within the turbopump that reduces axial shaft loads. Turbopump bearing life is determined by rolling-contact fatigue, which is a function of speed and load. ... The use of vacuum melted materials further increases life so the average predicted bearing life is approximately 65 times the B, life value.

A retractable seal for the propellant ball valves is a feature unique to the engine, added specifically to provide long life and reusability.

Automatic checkout, operational monitoring of flange leakages, and automatic propel!ant valve seat leakage detection have replaced the manual leak and functional techniques used on previous en­gine systems. Life monitoring techniques of in­ternal inspection, maintenance instrumentation, and drain system leak checks resulted from a maintain­ ability analysis conducted early in the design phase. Since corrective maintenance represents the largest single expenditure of resources during the SSME turnaround cycle, hardware accessibility and handling were emphasized early in the design phase. This maintenance concept results in a maintenance cycle for the three Orbiter main engines requiring an average of 25 hours of the 160-hour Orbiter turnaround.

a major effort was ex­pended to provide full internal inspection capabil­ity. This effort defined requirements, selected equipment, scheduled usage, and designed access ports.

The SSME was tested to at least 111% of rated power. For flight, 109% was the maximum.

Further reading:

This paper summarizes the changes made to the SSME over the life of the program. In several increments, they reduced the number of components and the time required to make them.

Another thing to look into is the RS-25E, the expendable variant of the SSME that is being planned for the SLS. Although as far as I know, they aren't planning any performance increases for that, just a decrease in manufacturing cost.

  • $\begingroup$ Since you have mentioned SSME. What about it in this case(only for SSME design and oppotunities that gives its design), if the engineers would decide to be single use, would they have chances to rise its performance (one of this elements the Isp,Thrust, Chamber Pressure). Could they have options to make a different "configuration" of this engine (its parts) to gain more performance? This answer would complete i guess the answer of my question. It looks difficult that i am asking you to answer yes or no, but it would be very helpful. $\endgroup$
    – Mark777
    Commented Aug 3, 2016 at 22:01
  • $\begingroup$ That's hard to say without a detailed understanding of the design and its margins. When a material approaches its limits, its behavior is not linear. Say you've got an engine that can run indefinitely at 1000 ºC. If you raise its operating temperature by 5%, engine life will not drop by 5%. It might still be good for indefinite operation at the new temperature, or it might melt after 30 seconds. $\endgroup$
    – Hobbes
    Commented Aug 4, 2016 at 12:54
  • $\begingroup$ Yes the expendable RS-25E as i said in to other comments above is a project focused more to make it low cost, so its primary objective is to be cheap not the high performance. As in the SSME case were perfomance was important but the primary objective was reusability. $\endgroup$
    – Mark777
    Commented Aug 4, 2016 at 16:34
  • $\begingroup$ It wouldn't necessarily need to be more complex, it just needs to survive re-entry heating. Or you can just reuse the whole stage like spaceX. If you want a reusable engine, as long as it is regeneratively cooled or film cooled or any kind of cooling that doesn't go away unless you shutdown the engine, will work and the engine will be reusable. $\endgroup$ Commented Apr 10 at 12:24

The dataset of rocket engines designed for reuse is possibly three or four?

  • The SSME as others have mentioned.
  • Merlin from SpaceX
  • Raptor from SpaceX
  • BE4 from Blue Origin (New Glenn and Vulcan)
  • BE-3 from Blue Origin
  • X-15's engine
  • RL-10 which was reused in the DC-X

Pretty much every other engine has been designed for expendability. But rarely in the true sense, which would be make it cheap and cluster them.

The SSME is considered very high performing and very expensive. Many will contest that it was even designed for reuse based on its actual history.

The information from the Merlin engine is that the same basic design has been used for both reuse and expendability. Beyond the need to be able to air start (Something the SSME lacks, and some initial Constellation designs considered) there does not seem to have been much that needed to be changed.

Obviously if you use an ablative nozzle design it does not make much sense in the reuse world.

In terms of performance, SpaceX has never indicated that there was any penalty for reuse of the Merlin engine. It was designed that way from the start. You want to be able to test your engine, again and again before flight, so at least some basic, marginal level of reuse is pretty much required.

The SSME were test fired before launches and then prepped for launch, more extensively then say a Merlin would require.

Some liquid boosters do hot fire testing, so even the expendable ones would need to be marginally reusable to accomplish this task.

So at some level, liquid fueled engines are usually built for marginal reusability. (Obviously solid engines, not so much).

Thus, it may be that the question is somewhat moot, since any engine that does hot fire testing, or acceptance testing of the engine, is likely close to reusable to start with, even if used expendable.

  • $\begingroup$ Yes but what about the performance is different or almost identical? $\endgroup$
    – Mark777
    Commented Apr 28, 2016 at 11:51
  • $\begingroup$ RL-10 is a borderline case you might want to add. It was originally designed as an expendable, but its straightforward and robust design gave it really impressive endurance, and the RL-10A-5 version was reused in flight testing of DC-X. $\endgroup$ Commented Apr 28, 2016 at 11:57
  • $\begingroup$ You also have to account for Raptor, Raptor Vacuum, BE4 and BE3U $\endgroup$ Commented Apr 10 at 12:25

I think you're asking the wrong question. Reusable engines will certainly have inferior performance to expendable ones, at least for a given cost/mass/size/TWR/ect. Just think about how cheap and simple model rocket motors are. They're nothing but explosives in a cardboard tube, and I can't imagine how you'd come up with a similar reusable system without making big sacrifices.

The real question of re usability is "How is a re-usable engine/vehicle recovered?" Consider how the SRBs from the space shuttle are recovered - by splashing down and being towed back. This makes landing simple - as all you need are parachutes to slow the fall - but introduces the problem of salt-water. Turns out that bathing your SRBs in a corrosive ocean is bad for them.

Now think of a SpaceX or Blue Origin style landing on land. It's much more complicated because of the need for landing gear, guidance, and the extra fuel to soft land, but I set my rocket down in relatively pristine condition. The engines aren't being eaten by salt water and don't need the same kind of intensive cleaning and refurbishment. That translates into less cost.

So, it's not so much about the engine as it is how you get it back.


I don't think this is a very useful question to ask. Many engines are more limited by the amount of money a company is willing to invest in both the development and the production cost of the engines. Both types of engines could be more performant if the builders were willing to invest a lot more money in r&d and/or in the engine production costs. If I interpret your question with an implicit "assuming r&d and production cost are similar" that would make it a lot more difficult to answer. The reasons for developing an engine and thus the budget are never the same.

Some examples:

  • As mentioned by @Russell the RS-68 was designed with an ablative nozzle coating instead of cooling the nozzle with fuel or making the nozzle of a more heat resistant material to reduce cost, even though this reduced the performance.

  • Nearly all existing US-built rocket engines until very recently used a gas-generator cycle, which is simpler but less performant than staged combustion cycles. The USSR did develop oxygen-rich staged combustion engines and those had a much higher performance than comparable US-built engines.

  • SpaceX is now the first company developing a full flow staged combustion engine (the Raptor), which is theoretically the most performant cycle. It is only because they are striving to make rockets reusable that it makes sense to invest in such an advanced cycle. With a non-reusable rocket it is cheaper to just make your engine and rocket a bit larger instead of investing all this money in a better cycle.

  • The only reusable system before SpaceX's rockets was the Space Shuttle, but its development budget was determined by political interests. Development cost was a major reason why it ended up with solid fuel boosters: high thrust, cheaper to develop, and 'reusable', but retrieving and refurbishing the SRBs was as expensive as just building new ones.

  • The US invested a lot in liquid hydrogen fuel rocket technology. This is more performant for high energy trajectories (i.e. interplanetary launches). For earth orbit launches there may also be some performance advantage depending on how you define 'performance', but it is not cheaper. The new space companies are not investing much in hydrogen, and Russia never built hydrogen fueled engines at all. But the ULA and ESA are continuing to use it because the technology now exists and they have experience in it.

If someone wanted to build the most performant (conventional) rocket engine, either reusable or not, where performance is measured by specific impulse and thrust, and money was no object, you would probably go for a full flow staged combustion hydrogen-oxygen engine, but no such engine nor plans for such an engine exist.

The definition of 'performance' is also not clear. If 'performance' just means thrust, you can always build larger engines of existing cycles, or a solid fueled rocket. If 'performance' just means specific impulse, you need to look at ion thrusters and other types of electrically powered engines, but those have very low thrust. The theoretical best would be an engine that emits photons for their momentum, i.e. a lamp.


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