My question is about operating costs of propellants, from production, handling, fueling, and anything else needed for the props. I am not just asking about how expensive it is to produce them. Operating costs must be taken into account, from production to transportation to loading and launching the rocket.

My meager understanding of these radically different propellants is giving me conflicting information:

  1. LH2/LOX is extremely cryogenic. It's hard to store and requires frequent replenishment. It's also more expensive to insulate everything from the tanks to the plumbing.

  2. UDMH/N2O4 is extremely acidic, toxic, sensitive to air, and apparently has just about every other terrible chemical property imaginable. It's hard to handle (transport and fuel up), requiring a lot more safety precautions (expenses).

So which is actually more expensive to use? LH2/LOX or UDMH/N2O4?

I'm sure there are some specific cases where one will be cheaper than the other, but in other specific cases it will be the other way around. Therefore, I want to be specific and list a concrete design problem:

Let's design a final rocket stage capable of providing a 3 km/s delta-v to a 10 ton payload. Would it be cheaper to use LH2/LOX or UDMH/N2O4? In other words, what is the cost to build and launch this stage using LH2/LOX, and then what is the cost to build and launch this stage using UDMH/N2O4?

And it doesn't need to be precise costs. I'm looking for which one will be relatively cheaper and why.

EDIT: since it seems that different nations will have different regulations affecting cost of these propellants, I will be clear and ask for the cost in the USA. However, if someone wants to do the same calculations/comparisons for Russia, that would be very useful too and certainly wouldn't go amiss.

This is admittedly not an easy question, but I believe it's answerable if you can find the right engineering data and/or experience (or white paper). I haven't been able to, however.

  • $\begingroup$ Do you also want to consider kerosene/LOX? $\endgroup$ Commented Jul 18, 2015 at 21:57
  • $\begingroup$ @RussellBorogove maybe. I left it out because I'm afraid it might be closed as too broad. My understanding is that Ker/Lox is more a middle-of-the-road compared to the two extremes of LH2 and hypergolics. I could be mistaken. $\endgroup$
    – DrZ214
    Commented Jul 18, 2015 at 22:16
  • $\begingroup$ Are you considering the cost to manufacture the rocket as well? Or just the operating cost of the fuel? $\endgroup$
    – PearsonArtPhoto
    Commented Jul 19, 2015 at 0:02
  • $\begingroup$ @PearsonArtPhoto the rocket as well. Or rather, the rocket stage. But the whole stage construction (and operation). I'll edit the bold question to make this clear. $\endgroup$
    – DrZ214
    Commented Jul 19, 2015 at 0:24
  • 2
    $\begingroup$ In my opinion you are asking someone to do a graduate thesis for you, or at least that level of effort. $\endgroup$ Commented Jul 20, 2015 at 1:27

2 Answers 2


For an upper stage, look to the Long March rocket series as a case history. For Long March 2-4, the boosters use hypergolic fuels while the upper stage burns LOX/LH2. Given that they have decades of experience with the hypergolic engines, that suggests LH2 is cheaper. Essentially they made the decision between building a bigger booster or adding strap-ons or going multi-core to increase payload, or else make the final stage lighter by using a higher ISP fuel. This also points out that you can't design a rocket in a (logical) vacuum. An upper stage capable of giving 3 km/s to a 10-ton payload may be cheaper to build, fuel, and operate using hypergolics, but the overall weight will be heavier than a LH2 fueled stage of the same performance, and that means boost stages must be larger and heavier and therefore cost more. A dollar spent on your upper stage may save five dollars getting it to altitude.

Also look to the J-2. Hypergolics are easy to restart. However the J-2 was restartable. Most importantly, the J-2 doesn't seem to have suffered from the same operational curses that boosters do when carrying LH2. From a purely empirical standpoint, if we flew the J-2 LOX/LH2 engines in Apollo upper stages without a lot of delays, there doesn't seem to be a compelling reason to spend more $ on a highly poisonous, heavier, corrosive replacement fuel.

That said, Long March 1 used a solid rocket upper stage. Hydrogen is tricky; development times are long and expensive.

Are you factoring system capital costs into it? R&D and building your hardware? Or is your question based on an existing sunk cost?

I am afraid that you are asking the same question that every rocket designer since Goddard has asked when they start a project with a clean sheet of paper. There is no definitive answer, only the answer which appears best, given all known factors and best modeling practices, for any specific design at a specific time.

Hypergolics are expensive and difficult to handle, but the engineering of a hypergolic engine is much simpler than for a LH2 cryogenic. So if you are looking for the fastest, cheapest system to develop and implement for a small number of launches, hypergolic is probably the better of the two. If you have more time and development money, and plan a longer service life for your system, you find yourself driven toward LOX/LH2. LOX doesn't actually appear that difficult to handle; it's the LH2 that kills you.

Per SF, $/kg of payload is your final engineering metric.

Look to history to inform your answer. Goddard and the V2 used LOX with gasoline and alcohol/water, respectively. The Titan 1 used LOX/RP1. For the Titan II, they modified the LR-87 engine into the hypergolic-fueled LR-87-5 so their ICBM could be stored with room-temperature fuel. So the decision was based on storage, not performance, and the engineering challenges were similar enough to modify a LOX/RP-1 engine rather than design something new. From this we can see that R&D and fabrication of a hypergolic engine is on par with that for LOX/RP-1 engines, which is about as cheap as liquid fueled rockets get. Hypergolic fuels are super expensive, but if your launcher has a short development cycle and limited R&D budget and you plan on a small number of launches, hypergolic wins. Actually, LOX/Kerosene wins, but that's not your question.

If you have 30 years and billions of $ to iterate your design, then LOX/LH2 wins. The proof is the Delta IV and its RS-68. If decades of engineering experience pointed to a hypergolic booster as getting payloads up more cheaply (per kg of payload to orbit), ULA would be putting money into hypergolic, or pushing the government to fund a new development effort.

I have a bias. I hate LOX/LH2 systems. LH2 is simply evil. It seeps through "cracks" in welds which any other material would consider perfectly impermeable. Hot hydrogen makes metal BLISTER. It's so cold that the foam insulation on the shuttle tanks had to be foamed with helium; foaming with air leads to the air condensing and the foam collapsing. I feel that if the shuttle program had had fewer delays due to tracking tiny hydrogen leaks, they may have been more willing to address real concerns like the SRB O-rings. I consider it an engineering miracle that they have managed to "tame" LH2 and launch the Delta IV's on schedule. Considering they are building on SSME technology, it's a miracle about 45 years in the making. Also, note that ULA only uses the Delta IV when it can't fit the payload onto a LOX/RP-1 fueled Titan, and the Delta is planned for phase-out once they get a LOX/Methane booster working.

And that's why LOX/RP-1 has regained popularity, especially in boosters. The lower ISP doesn't hurt performance nearly as badly as it does in an upper stage. Sure, it's "1950's technology", but as such it has 70 years of engineering refinement and leads to a $/kg of payload much lower than for competing LH2 systems.

Given my bias, my answer is "Neither." For a first stage, unless you have a nearly unlimited development budget and schedule, go with LOX/RP-1 or LOX/LMethane for your cheapest $/kg payload. That seems to hold true for the smallest to the largest launch systems.

Second stage? More engineering decisions, but LH2 is probably your winner. Look to the J-2 as your case history. The poor ISP of hypergolics will hurt your overall system performance more than on a first stage. Hypergolics are easy to restart. However the J-2 was restartable. Most importantly, the J-2 doesn't seem to have suffered from the same operational curses that boosters do when carrying LH2. From a purely empirical standpoint, if we flew the J-2 LOX/LH2 engines in Apollo upper stages without a lot of delays, there doesn't seem to be a compelling reason to spend more $ on a highly poisonous, heavier, corrosive replacement fuel.

And have you read up on Nitrogen Tetroxide? That stuff is evil. As far as I can tell, if you can smell it, then you are going to die.

  • $\begingroup$ To the best of my knowledge, we still don't have much for non-cyrogenic orbital maneuvering thrusters though. I wasn't aware N2O4 was quite that bad -- I associated it with basically "gaseous nitric acid" -- though I certainly have never smelled it despite having once upon a time made my own WFNA. $\endgroup$
    – ikrase
    Commented Dec 6, 2019 at 9:50
  • $\begingroup$ The apollo-soyuz crew managed to suck a lot of N2O4 into the capsule during the descent, when they failed to deactivate the rcs before opening the cabin atmosphere valves. They reported smelling it (strongly), seeing brown vapors, and had difficulties breathing (CMP passed out). Burning eyes and respiratory issues for a few days, not more. The medical report is available and covers it in good detail. "smell it and you die" is quite an exaggeration. $\endgroup$
    – Innovine
    Commented Apr 16, 2022 at 8:08

Operating costs must be taken into account, from production to transportation to loading and launching the rocket.

...and since the question is about application in the final rocket stage, that must include cost of getting the final stage to altitude and speed where that stage is ignited.

And here the comparison to cryofuels crashes and burns.

UDMH/$N_2O_4$ has a poor density impulse, 316-kg s/l. LOX/LH2 - 124 kg-s/l.

That means you need to load up much more of it onto the final stage, than say, LOX/LH2 to achieve the same delta-V.

And that means, that regardless of costs of production, transportation, loading and fueling up the final stage, the cost of about the most expensive part of the rocket - the launch stage - goes up by strides. Any benefits to reduction of cost or complexity of the final stage will be completely overshadowed by the increased cost of the initial stages that need to handle the increased payload.

  • $\begingroup$ The density impulse numbers you cited seem to disagree with you. 316 is 2.5x more than 124. That suggests we would need much more LOX/LH2 instead of the other way around. Did you typo a number? $\endgroup$
    – DrZ214
    Commented Feb 2, 2016 at 0:10
  • $\begingroup$ @DrZ214: I used the same chart that Paul had linked in his answer, and it appears unlike Specific Impulse, "more is worse" - Hydrazine is 439, Solid fuels are 474, and so on. These are known to have a very poor specific impulse. $\endgroup$
    – SF.
    Commented Feb 2, 2016 at 8:46
  • 2
    $\begingroup$ The primary metric for rocket propellant is specific impulse where more is better. If you compare the Isp for the Atlas and Titan rockets, you will see it is very similar. You have also misunderstood density impulse in another important way: higher is better as it means a smaller tank can produce the same net change in velocity. In other words, the low density impulse of LOX/LH2 is bad because it means bigger tanks are required. $\endgroup$
    – user8269
    Commented Mar 14, 2016 at 17:20
  • 1
    $\begingroup$ @PhilipNgai: Specific Impulse is a nice engineering metric that can serve as a good rule-of-thumb which propellant is better, but the ultimate metric is still (delta-V)/\$ (where \$ includes both cost of the engine, the propellant, and getting them to where they operate). That's why most satellites still use monopropellant engines for RCS and maneuvering instead of ion - because for the low needed delta-V, the crappy ISp of monoprop engines still beats the superior ion engines in terms of cost of construction and is similar in costs of delivery. $\endgroup$
    – SF.
    Commented Mar 14, 2016 at 20:39
  • $\begingroup$ Usually better ISp means less mass for the same delta-V, and that means lower cost, but sometimes costs of the engine itself just outweigh the fuel mass savings - and so, a technology of worse ISp is used because it costs less. Why isn't Soyuz loaded up with a big array of ion drives instead of chemical rockets? $\endgroup$
    – SF.
    Commented Mar 14, 2016 at 20:47

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