Okay, I admit it: I’m a sucker for the Olympics. I watch with rapt attention to sporting events that I would otherwise never consider viewing other than under the once-every-four-years heading of the Olympics. Why is that? Perhaps that is somehow a measure of my shallowness as a sports fan. Nevertheless, I was truly on the edge of my seat watching the women’s team archery semi-finals and finals. Great drama. Wonderful competitors. Exceptional skills. Bravo ladies!
Another thing that I find fascinating about the Olympics is the fact that it brings together such a broad range of people. No, I’m not going to trail off in a chorus of Kumbayah. You simply cannot deny, however, that during the opening ceremonies you see people of every possible color and shade, from every corner of the planet, straight hair, curly hair, black hair, blonde hair, red hair, eye colors to fill a rainbow, and most startlingly, such an amazing collection of body types. These are all world-class athletes and yet they’re often so different from each other. I like seeing the six-foot-seven volleyball player walking next to the four-foot-ten gymnast. I like seeing the contrast of the marathoner and the shot-putter. We’re all the same species, but, my goodness, we come in an amazing array of shapes and sizes and various accoutrements.
The Pivot to Topic
Rocket engines, too, come in an array of shapes and sizes and various accoutrements (…I bet that you were wondering when or how I’d turn the conversation on topic). I know that this is a blog dedicated nominally to J-2X development, but I think that it’s important to understand where the J-2X fits in this family of rocket engines. So, let’s start with a table of top-level engine parameters:
Note that is list is nowhere close to being comprehensive. There are lots and lots of rocket engines out there including those currently in development or in production and many that have been retired (like the F-1A in the table). And if you open the window a little wider to include engines originating from beyond our shores, then you’ve got many more Soviet/Russian, European, Japanese, and Chinese engines to consider. All I want to do here is expose you to some basic yet significant differences between this small set of examples. Interestingly, if you can understand these few engines, then you can understand most of rest of the ones out there as variations on these basic themes.
Please allow me to introduce you to the engines listed in the table.
• Of course, the J-2X needs no further explanation for anyone who reads this blog regularly.
• The RL10 is a small engine that has been the product of Pratt & Whitney since the late 1950’s. Over the past sixty years it’s evolved and matured. It was actually used on a NASA vehicle back in the 1960’s, the Saturn I launch vehicle upper stage (S-IV). Today it’s used, in different variants, as an upper stage and in-space engine for both the Atlas V and Delta IV launch vehicles.
• The RS-25 is another name for the Space Shuttle Main Engine (SSME). The development of the SSME began with research efforts in the late 1960’s, using a great deal of knowledge gathered from the development of the original J-2, and it was first tested in 1975 and first flew on STS-1 in 1981. The RS-25 engine is now designated to be the core stage engine for the next generation of launch vehicles under the Space Launch System (SLS) Program.
• The F-1A was an upgraded version of the F-1 engine that powered the first stage (S-IC) of the mighty Saturn V launch vehicle that first took man to the Moon. The F-1A was a more powerful version of the F-1 with a handful of design changes intended to make it cheaper yet more operable and safe.
The Key is in the Power
In a blog article here over a year and a half ago, I introduced you to the gas generator cycle engine. The key philosophical point discussed in that article about what makes a rocket engine an engine is the fact that it feeds and runs itself. It does this by finding a means for providing power to the pumps that move the propellants. The origin for this power is the key to any rocket engine cycle. In a gas generator engine, this power is generated by having a separate little burner that makes high-temperature gases to run turbines that makes the pumps work. Below is a schematic for such a system. You’ve seen this schematic before and it is very much like J-2X.
MCC = Main Combustion Chamber
GG = Gas Generator
MFV = Main Fuel Valve
MOV = Main Oxidizer Valve
GGFV = Gas Generator Fuel Valve
GGOV = Gas Generator Oxidizer Valve
OTBV = Oxidizer Turbine Bypass Valve
Behold Now Behemoth
The F-1A power cycle is similar to the gas generator cycle shown above in that it is still a gas generator cycle, but rather than two separate turbopump units, there was only a single (huge) unit that contained both pumps. So, a single turbine was used to power both pumps rather than having two separate turbines like J-2X. Going back to the table, you will see that the F-1A was different from the J-2X also in the fact that the propellants were different. The J-2X uses hydrogen for fuel and the F-1A used RP-1 (FYI, RP-1 stands for “rocket propellant #1” and is actually just highly purified, high quality kerosene). The chief difference between hydrogen and kerosene is chemistry. A hydrogen-fuel engine will get higher specific impulse than a kerosene-fuel engine but kerosene engines have the distinct advantage of being able to generate more thrust for a given engine size. With a kerosene engine, you are simply throwing overboard more massive, high-velocity propellants in the form of combustion products. Hydrogen is light and efficient from a “gas mileage” perspective but kerosene gets you lots and lots of oomph. That’s why you typically use it for a first stage application like on the Saturn V vehicle. You want to have lots of oomph to get off the ground. Later, on the upper stages, you can better use the greater gas mileage afforded by hydrogen.
Note, however, that you could theoretically build a hydrogen engine as large as the F-1A in terms of thrust. The RS-68 (also a gas generator cycle engine) on the Delta IV vehicle puts out around three quarters of a million pounds-force thrust so that’s pretty big. Also, back in the 1960’s, there was conceptual design work performed on an enormous hydrogen fuel, gas generator cycle engine called the M-1. On paper, that behemoth put out 1.5 million pounds-force of thrust just like the F-1 on the Saturn V. But that project was abandoned and here’s why: hydrogen is very, very light so if you want to carry any appreciable amount, you need to have truly huge tanks. Huge tanks mean huge stages. Huge means heavy. Eventually it becomes a game of diminishing returns at the vehicle level.
What this discussion of J-2X and F-1A (and RS-68 and even M-1) shows you is the extreme versatility of the gas generator cycle. It can be used with nearly any reasonable propellant combination and it can be scaled from pretty darn small to absolutely enormous.
Shaving with Occam’s Razor
Occam’s Razor is the notion that one should proceed with simplicity until greater complexity is necessary. Along these lines, I will introduce you to a simpler engine cycle: the expander cycle. For this engine cycle, you do not use a gas generator to drive your turbine(s) so you don’t have a second, separate combustion zone apart from the main combustion chamber. That makes everything simpler. Instead, you use only the heat gathered in the cooling the thrust chamber assembly (i.e., the main combustion chamber walls and that portion of the nozzle regeneratively cooled). See the schematic below.
See? I got rid of not just the gas generator but also the two valves that fed the gas generator. That’s huge in terms of simplification. And whenever you can make an engine simpler you’ve usually made it cheaper and more reliable just because you have fewer things to build and fewer things that could break. Cool!
Here, however, is the problem: How much power do you really have just from the fluid cooling the walls? The answer can be found by looking at the table and seeing, for example, the RL10 thrust output is less than one-tenth of J-2X. You just can’t pull that much energy through the walls. There have been attempts to increase heat transfer by various means including making the main combustion chamber longer than typical so that you have more heat transfer area or even by adding nubs or ridges onto the wall to gather up more heat. Using the longer chamber notion, the European Space Agency is working on an engine called the Vinci that almost doubles the thrust output from the RL10, but getting much further beyond that is darn tough. Also note that hydrogen is a wonderful coolant based upon its thermodynamic properties. Being a wonderful coolant means that it picks up a lot of heat. It is difficult to imagine using the expander cycle engine with another fuel beside hydrogen (though maybe methane might work … haven’t examined it).
On the plus side, in addition to the simplicity, what the cycle shown offers is what is called a “closed cycle” meaning that no propellants are thrown overboard other than through the main injector. In a gas generator cycle engine, after the gas generator combustion gases pass through the turbine(s), it’s dumped into the nozzle (or, in other schemes, dumped overboard in other ways). Any propellants or combustion products that do not exit the rocket engine through the main injector and through the main combustion chamber throat represent an intrinsic loss in performance. “But,” you’ll say, “the specific impulse for the RL10 and the J-2X in the table are the same.” Well, that’s a little bit of apples and oranges because it’s based upon the nozzle expansion ratio. Another model of the RL10, the B-2, has a much larger nozzle extension and the vacuum specific impulse for that model is over 462 seconds (minimum). The European Vinci engine that I mentioned above has a projected vacuum specific impulse of about 465 seconds. Those are darn impressive numbers that make the mouths of in-space stage and mission designers drool.
A couple of final notes about the expander cycle engine. First, the RL10 is not quite like the schematic shown. It only has one turbine with one pump driven directly and the other pump driven through a gear box. Thus, the OTBV goes away (making it even simpler!). Second, there are versions of the expander cycle engine concept that are not closed cycles. In these versions, you dump the turbine drive gas overboard in a manner similar to what you do in a gas generator cycle. You are still using the heat from the chamber walls to drive the turbine(s), so it’s still an expander, but with an overboard dump you can also leverage a larger pressure ratio across the turbine(s) and thereby get a bit more oomph out of the cycle. You sacrifice a bit of performance for more oomph. The Japanese LE-5B engine is an open expander cycle engine like this (also called an “expander bleed” cycle).
“We do these things not because they are easy…”
So, you’ve seen the incredibly versatile gas generator cycle engine. And, you’ve seen the simple yet limited expander cycle engine. So what do you do if you say, “The heck with it, I want the Corvette”? What if you want a closed cycle, high performance engine not limited to lower thrust levels and you’re willing to accept consequent greater complexity? The answer is staged combustion. Below is a simplistic schematic for a staged-combustion engine.
CCV = Coolant-Control Valve
PBOV = Preburner Oxidizer Valve
In a staged combustion cycle engine, we rename the gas generator and call it the “preburner.” The biggest difference between a gas generator cycle and a staged combustion cycle is what you do with the turbine exhaust gases. In a gas generator cycle, the turbine exhaust gases effectively get dumped overboard. In a staged combustion cycle, the turbine exhaust gases get fed back into the main injector and get “burned again.” This is possible since the combustion in the preburner is off from stoichiometric conditions, meaning that in addition to combustion products you also have lots of leftover propellant (either fuel or oxidizer depending on the scheme). The leftover propellants from the turbine exhaust then become part of the mix of propellants in the main combustion chamber.
That sounds simple, right? It’s just a twist on the gas generator cycle theme, right? Well, there are larger implications. First, think about the pressure drops through the system. On a gas generator cycle engine, the pressure in the gas generator can be lower than the main chamber. After all, the downstream side of the turbine(s) is effectively ambient, external conditions. In a staged combustion cycle, the preburner pressure has to be substantially higher than the main chamber pressure sitting downstream of the turbine(s) or you don’t get enough flow to power the turbine(s). Insufficient turbine power and the cycle doesn’t work. So, in general, a staged-combustion cycle engine has higher system pressures than a gas-generator cycle engine of comparable size. Next, think about starting the system. In a gas generator cycle engine, the two combustion zones are effectively disconnected. In a staged combustion cycle engine, the two combustion zones are on either side of the turbine(s) so there is effectively communication between these two zones. Now, try to imagine getting these two combustion zones ignited and up to pressure and the turbine(s) spun up to speed in an orchestrated manner during the start sequence. It ain’t easy.
So, what do you get for this complexity and higher operating conditions? Well, you get a closed cycle, high performance, and high thrust engine design choice. The RS-25 (SSME) is the American example of such an engine. If you put a higher expansion ratio nozzle on the RS-25, just as with the RL10 discussion, the specific impulse value would be as much as ten seconds higher than J-2X. However, if you go out and find a schematic of an SSME, what you’ll see is a heck of a lot more complexity than even I’ve shown in my simplified sketch. Because the pressures are so high, there are actually four separate turbopumps and a boost pump in the SSME. The design relies on putting pumps in series to achieve the necessary pressures and fluid flow rates through system. And, the SSME has not one but two separate preburners, one for the high pressure fuel turbopump and one for the high pressure oxidizer turbopump. It’s a very complex engine, but it has extraordinary capabilities.
The RS-25 (SSME) is a staged combustion cycle engine with hydrogen as the fuel. The preburners are run fuel-rich such that the generated gases contain excess hydrogen for injection in the main chamber. Back in the days of the Soviet Union, they developed a whole series of staged combustion cycle engines that instead used kerosene as the fuel. In these engines, the preburner is run oxidizer-rich so that the gases run through the turbines and then through the main injector have excess oxidizer to be used for final combustion in the chamber. The Russian-supplied RD-180 that is currently used for the Atlas V launch vehicle is an example of such an engine. It too is an extremely complex, high pressure, and high performance engine.
So, staged combustion cycle engines are not easy. Their complexity and operating conditions suggest, generically, greater expense and lower reliability. But if you can make the trade-off between high performance and the adverse issues, then they can function quite impressively. Nearly thirty years of Space Shuttle flights are an indisputable demonstration of this fact.
Just One Bolt
Can you imagine opening a hardware store and selling just one kind of bolt? That would be it. One brand. One diameter. One length. And just one bin full of identical versions of this one bolt in your store. It sounds really kind of stupid. The unavoidable truth is that you need different bolts for different applications. It’s kind of like trying to imagine telling the Olympic gymnastics team that they now had to play basketball and the basketball players to do gymnastics. I don’t know about you, but I’d love to see Lebron James have a go at the pommel horse.
Well, over the last fifty-plus years, we’ve developed different rocket engines and rocket engine concepts for a variety of different applications. Just one design does not fit all applications. Each design has advantages and disadvantages. If you can understand the basics of what I’ve discussed in this article, however, then you will have a fundamental understanding of at least 90% of the engines spanning that fifty-plus years of history. And that, in turn, might help you better appreciate why one bolt is chosen over another or why, for example, shot-putters tend to be a bit more beefy than cyclists.
6 thoughts on “Inside The J-2X Doghouse: Beyond the Gas Generator Cycle”
I share your fascination with the Olympics. For me it is the degree of preparation and sacrifice in a high risk endeavor. Where the reward is not monetary just the satisfaction of having done your personal best, and possibly (low probability) earned a few seconds of glory before a worldwide audience.
Anyway, back to rockets. As a systems hardware designer (not in the aerospace field) I have long been fascinated with rocket development. I just came across a gem in the field: Mulready’s “Advanced Engine Development at Pratt & Whitney”. I suspect you must be familiar with this book. He devotes a chapter to the RL10.
Also, I recently came across your blog. You have a wealth of information here that is certainly worthy of a book where Mulready left off (1971).
Looking forward to reading more of your blog as I peruse Dick Mulready’s book.
Aside from the kerosene vs hydrogen and the preburners running oxidizer rich vs fuel rich what are the other major design differences between the RD-180 and the RS-25?
I am familiar somewhat with the history of the RD-180 being derived from the quad RD-170. You could probably give us a pretty good insight into some of the nuances and design details. The RD-180, so far, has a pretty good track record with the Atlas family.
Also, how about adding the RD-180 to your table above?
@Tariq: Sorry for the delay in the response. August has been a terribly busy month for me.
With regards to RS-170/180, I am not an expert. I have a schematic of RD-170 sitting right in front of me, but I’ve never actually worked directly with these engines. I do know some folks around here who have worked with the RD-170 in support of the Atlas V vehicle development effort some years ago, but because this is a Russian engine, they tend to hold close to the vest some of the more interesting details. This is entirely understandable from both a proprietary data perspective and in terms of protecting their national security.
Here’s what I do know about them. The RD-170 was developed originally to power the boosters for the Soviet version of the Space Shuttle, the Buran. It later became the mainstay of the Zenit launch vehicle. It is a liquid oxygen (LOX) / kerosene engine with a mammoth thrust level (approximately 1.8 million pounds-force). It basically puts out the same thrust level as the F-1A but does so with a greater efficiency (337 seconds vacuum specific impulse versus 269 seconds for F-1A). That’s because it is a staged-combustion engine rather than the gas-generator cycle of the F-1A.
The Soviets did a lot of work with oxygen-rich staged combustion engines. In addition to the RD-170/171/180 family, there is also the NK-33 (originally developed for the Soviet moon program and today being reincarnated through Aerojet as the AJ36). In this country, oxygen-rich staged combustion is not something that we’ve pursued (so far as I know) beyond conceptual design and laboratory experiment and/or sub-scale component testing. Working with a hot, gaseous, oxygen-rich environment is darn scary. It requires some unique material considerations to say the least. The Soviets/Russians invested enough effort to get comfortable with this approach. We’ve have not.
A somewhat unique aspect of the RD-170 in terms of why it looks so different than an F-1A is the fact that it has four separate combustion chambers. Back when the F-1 was being developed in the United States back in the 1950’s and early 1960’s, one of the biggest recurring issues was combustion instability. Kerosene engines are notorious for this. Simply put, kerosene is a big molecule. Busting it up during the combustion process takes some time. No, not much, but some. If that characteristic time (dependent upon injection details) links up with a characteristic mode of the chamber (determined by geometry and the speed of sound of the gases in the chamber), then you could inadvertently wander into a situation where any minute disturbance and grow out of control and destroy the whole rocket. In the U.S., we solved this issue with a complex baffle arrangement on the main injector (look up the throat of an F-1 should you get the chance and you’d see what I mean). Despite the difference in cycle, the Soviets also had issues with combustion instability. One aspect of their solution involved going to multiple, smaller combustion chambers. Thus, where the F-1 has one huge turbopump feeding one huge thrust chamber, the RD-170 has one huge turbopump feeding four smaller (yet still pretty big) thrust chambers.
The RD-180 is, effectively, half of an RD-170. Rather than four thrust chambers, it has two. The thrust level is therefore half of the larger engine but the efficiency of the staged-combustion cycle is retained.
Bill: Thanks for the details, especially your comment about the kerosene molecule.
Are you aware of any new staged combustion booster rocket designs in the US since the RS-25? The only one I am aware of is the Spacex Merlin-1D. They seem to have learned from the lessons of the past by going to a single thrust chamber per engine and using lots of them ( 9 in Falcon9 and 27 in Falcon Heavy!). I wonder if the -1D runs oxidizer rich? Clearly, they are keeping most of the details close to the chest. Comments?
I believe that the Merlin 1-D is a Lox/RP-1 gas generator cycle at a little more than half of the thrust level of J-2X. Whether SpaceX is working on a staged-combustion version of their Merlin series, I don’t know. That would be quite interesting if they were.
I don’t know of anyone (other than the Russians, of course) who is working with ox-rich, Lox/RP-1, staged combustion cycle engines other than on-going research activities.
Bill: I knew that the current Merlin1 used the gas generator cycle, but I could have sworn that I heard Elon Musk announce that the 1D was going to be a staged combustion engine. Looking at the closeup of the 1D on the test stand it is clear that it is indeed a gas generator. You can see the dark exhaust fumes spewing out next to the main chamber. No details on Merlin 2 yet.
So, clearly the Russians are the only game in town on ox-rich closed cycle engines for the moment.
I am curious as to issues with using ethanol instead of RP-1. I know people are using lox/methyl alcohol.
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