Monthly Archives: August 2012

J-2X Extra: What's in a Name?

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It’s been over six years since I started working on the J-2X development effort.  I missed the very first day that the notion of a J-2X engine was conceived, but I was only two weeks late to the party.  So, I’ve been with the thing almost from the beginning.  And throughout that entire period, whenever I get the chance to talk to people outside of our small, internal rocket engine community (…for very understandable reasons, they don’t let us out much), the single, most frequent, recurring, and ubiquitous question that I hear is something along the lines of this:

“How come you guys are spending so much time and effort recreating an engine that flew nearly fifty years ago?”

That is an entirely fair question.  I am not a volunteer.  As generous and as charitable as I like to consider myself, I do accept a paycheck.  So do my coworkers.  So does our contractor.  Thus, all this work to develop J-2X isn’t free and, as I said, the question asked is therefore a valid point of discussion.

To a certain degree, I tried to answer this question by way of analogy in a J-2X Development Blog article posted a year and a half ago (December 2010) about a 1937 Ford Pickup truck.  But analogies and metaphors can sometimes be abstruse.  Let us eschew obfuscation and arrive expeditiously to the point:  What makes J-2X different from J-2?

The J-2 rocket engine, developed by Rocketdyne and the NASA Marshall Space Flight Center, was qualified for flight in 1966.  Between August 1966 and January 1970, 152 engines were produced.  Between 1962 and 1971, some 3,000 engine tests were conducted.  The J-2 engines were used for the second stage of the Saturn 1B vehicle and the second and third stages of the Saturn V vehicle.  (Note that I wasn’t much involved in the original J-2 project considering that it was concluding just as I was figuring out that whole reading thing in First Grade.  Remember Dick and Jane, Sally and Spot?)

The most significant differences between these two engines can be found in their performance requirements.  I suggest that these are most significant because it is these differences that lead directly to a majority of the physical design differences between these two engines.

That’s an increase in thrust level of over 25%.  And the specific impulse increase is on the order of 6%.  While that doesn’t sound like much, in the realm of rocket engines, given that the J-2 and the J-2X are using the same power cycle, it’s huge.  It means that we’re pulling staged-combustion or expander cycle levels of performance from a gas generator engine.  That’s really something special.

From requirements flows form.  Or, as stated by architect Louis Sullivan (mentor to Frank Lloyd Wright): “Form follows function.”  You don’t design and build a rocket engine a certain way because it’s neato.  It’s designed to meet requirements that fulfill mission objectives.  It’s not like a 1959 Cadillac where stuff was added just because it looked really cool (picture below courtesy of the Antique Automobile Club of America Museum in Hershey, PA).

In order to get that kind of boost in performance for J-2X, we had to do two fundamental things:  (1) move more propellant mass through the engine, and (2) use that propellant more efficiently.  To the first point, J-2X pumps into itself and expels out approximately 20% more propellant per second than did J-2.  That translates into needing a whole lot more pumping power.  Here’s a comparison of power requirements for the J-2 and J-2X pumps (as a point of reference, a typical NASCAR engine generates about 750 horsepower):

That’s between 80% and 90% more power for J-2X as compared to J-2.  The reason that you need so much more is not only the need for greater flow, but also the need for more efficiency in usage is manifested as higher discharge pressures.  I’ll explain this further below.  But first, let’s talk about the hydrogen pump just because it’s an interesting story.

Back in the day, when J-2 was first being conceived of, the technology of how exactly to pump liquid hydrogen was still being developed.  The RL10 engine existed already, but it was about 1/10th the size of J-2.  Some work had been done with pumping hydrogen as part of the NERVA nuclear thermal propulsion development effort, but not everything learned there was widely distributed.  This relative lack of information resulted in J-2 having a liquid hydrogen pump that was, in reality, an axial compressor.  You see, the problem is that liquid hydrogen is so light that it kinda sorta acts as much like a gas as a liquid.  I’ve heard it described as being like whipped cream but less sticky. 

So, do you pump it like a liquid or like a gas?  You typically use axial compressors for gases.  That’s what you use in turbojets for airplanes.  And you can get it to work with liquid hydrogen, as J-2 clearly demonstrated, but it’s not the best solution.  One of the issues is that a compressor has some unfortunate stall characteristics where the effectiveness of the pump can plummet during the start transient.  This is caused by what is known as the start oscillation that always happens in liquid hydrogen engines.  Picture this:  Prior to start, everything up to the valves that hold back the flow on the hydrogen side is chilled down to liquid temperatures (typically 36 to 39 degrees Fahrenheit above absolute zero).  Then the valves open during start sequence and the liquid hydrogen suddenly comes into contact with relatively warm downstream metal.  The result is similar to what happens if you sprinkle water into a hot frying pan.  In other words, the liquid boils immediately upon contact.  In a rocket engine this causes a transient “blockage” as this voluminous plug of newly formed hydrogen gas gets pushed through the system.  In terms of the pump, this sudden “plug” downstream results in a transient, elevated pressure at the pump discharge and this can cause the pump to stall, especially if it’s an axial compressor.  In order to overcome this effect, they had to precede the J-2 start sequence with several seconds of dumping of liquid hydrogen through the whole system to pre-chill the metal downstream of the valves. 

Okay, so that’s not too much of a big deal, but it was a nuisance.  By the end of the 1960’s, it was clear to most folks that the better way to pump liquid hydrogen was to use a centrifugal pump and that’s the way we’ve done it ever since (including on the J-2S engine, which was an experimental engine tested in the early 1970s as a follow-on to J-2).  With a centrifugal pumps, you get to avoid the stall issues inherent with an axial compressor and you get a more compact, powerful machine.  Which is good considering how much more power we need to pull out of the pump for J-2X.

In addition to changing from axial to centrifugal, we had to make a number of other changes to the turbomachinery.  In one place, we used to use on J-2 an Aluminum-Beryllium alloy.  Well, you can’t use Beryllium anymore since it is considered too dangerous for the machinists working with the metal on the shop floor.  In particular, Beryllium dust is toxic.  And since we really like the guys working on the shop floor (as well as following the law), we had to go to another alloy.  Also, we redesigned internal seal packages and rotor bearing supports using the most modern analysis and design tools and methods.  In short, there’s not much in the turbomachinery, both fuel and oxidizer, that wasn’t reconsidered and redesigned to meet the imposed requirements.

Now, the other reason that we need 80% to 90% in addition to pumping 20% more “stuff,” is the fact that we had to get that stuff to higher pressures.  Why?  As discussed in a recent previous blog article, if we go to a higher combustion chamber pressure, then we can have a smaller throat and, with a smaller throat, we can have a larger expansion ratio without getting too out of hand with engine size.  And, because of our extreme specific impulse requirement (remember: form follows function), we need that very large expansion ratio.  So here are the top-level thrust chamber parameters:

The J-2 main combustion chamber was built from an array of tubes braze-welded together.  When you needed the walls of that chamber to be actively cooled, this was the most common way to make combustion chambers “back in the day.”  This is a fine method of construction, but it is kind of limited in terms of how much pressure it can contain.  For the Space Shuttle Main Engine project in the early 1970’s, we needed the capability to handle a much higher chamber pressure and so we (i.e., Rocketdyne working in coordination with NASA) developed what is called a “channel-wall” construction method.  So, to get the higher performance using the higher chamber pressure, we had to abandon the tube-wall construction method for the J-2X main combustion chamber and use a channel-wall main combustion chamber similar to the Space Shuttle Main Engine.

The main combustion chamber is on the top end of the scheme to get the larger expansion ratio.  On the bottom end, we had to add a large nozzle extension.  On the J-2, the nozzle consisted of another tube-wall construction.  For J-2X, we have a tube-wall section that is actively cooled and then we have the radiation-cooled nozzle extension beyond that.  The reason for transitioning is because the nozzle going out to a 92:1 expansion ratio has a diameter of nearly 10 feet and a tube-wall construction that large would be unreasonable heavy.  In other words, from the vehicle perspective, the engine would be so heavy that its weight would offset any benefit from performance.  The radiation-cooled nozzle extension is significantly lighter.

That make it sound easy, doesn’t it?  If you want more performance, just strap on a big hunk of sheet metal and call it a nozzle extension.  I wish that it were that easy.  First, you need to figure out what material to use.  Metal?  Or maybe carbon composite?  Plusses and minuses for both.  Then you need to learn how to fabricate the thing light enough to be useful.  And then you have to make it tough enough to survive the structural and thermal operating environments.  In the pictures immediately above you can see a sample panel of how the J-2X nozzle extension is made and you can also see one of these samples sitting in a test facility where we blasted the panel with high velocity hot gases to partially simulate nozzle flow environments.  The panel has a coating that enhances the radiation cooling so not only does the panel itself have to survive the environment, but so does the special coating.

Other things that we’re doing to get more performance out of the engine include the use of a higher density main injector and the use of supersonic injection of the turbine exhaust gases into the nozzle.  When you talk about “injector density,” what you’re talking about is the number of individual injectors stuffed into a given space.  Up to a point, the more injectors that you have, the better mixing you get, and, from that, the better performance you an extract from the combustion process.  The picture below shows some testing that was done early on in the J-2X development effort to optimize the main injector density.

With regards to the turbine exhaust gas, on J-2 it was effectively dumped into the nozzle with the only intent being to not mess up the primary flow.  For J-2X, we carefully designed the exhaust manifold and internal flow paths to get as even a distribution as possible around the nozzle and, from there, we are injecting it into the flow through mini throats at supersonic velocity.  Here again we are extracting as much performance as we can given the simplicity of the power cycle.

The next element of the engine to consider is the thing that creates the power that drives the turbines…that spins the pumps…that feeds the injectors…that fill the chamber…that makes thrust.  In other words, I’m talking about the gas generator.

So, due to the increased power needs of the pumps, the gas generator has to flow twice as much propellant and at higher pressures through the turbines as compared to J-2.  The temperatures are pretty much the same since this parameter is mostly limited by material properties of the spinning turbine components.  In terms of “form following function” from a design and development perspective, these increased power requirements translated to the fact that gas generator used for J-2 was entirely inappropriate for J-2X.  It just wouldn’t work.  Rocketdyne had to design a new gas generator based upon work that they had done as part of the development of the RS-68 rocket engine (used on the Delta IV vehicle).  In the past, I’ve shown some pictures and even video of the whole development test series that we conducted to validate the design of our gas generator.  Below is a representative picture of our gas generator component test bed. 

Something not captured in the table of performance requirements way up above is the bevy of requirements imposed on the J-2X in terms of health monitoring and controls functionality.  These too resulted in differences between J-2 and J-2X. 

The J-2 engine had a sequencer to control the engine.  Yes, it consisted of solid-state electronics, but other than that it was pretty much like the timer on your washing machine.  The J-2X has an engine controller, which is a computer with embedded firmware and software that allows for a great deal of functionality in terms of engine control and system diagnostics.  Some of these diagnostics we call redlines.  These are specific limits that we place of measured parameters such that, should we break the limit, then we know that something bad has happened to the engine.  The idea is to catch something bad before it turns into something potentially catastrophic.  This is all part of the higher reliability and safety standards that have been applied to J-2X as compared to J-2.

The J-2X controller is composed of two independent channels such that if one fails, the other can take over.  For critical measurements that inform the controller during engine operation, we actually take four separate measurements, compare them to make sure that they’re reasonable and good, and then use algorithms to perform the health checks.  That’s one result of the imposition of more detailed requirements pertaining to reliability and safety.  Along these same lines, we also have a number of design, construction, and workmanship standards that were applied to every aspect of the J-2X engine design, development, and fabrication.  These standards, in combination with more evolved and advanced analysis tools, have, in a number of cases, further driven design changes away from heritage J-2 designs to what we’d call modern human-rated spaceflight hardware.

In an old J-2 manual, I found reference to a reliability value for that engine equivalent to 2,000 failures per one million missions.  The requirement for J-2X is 800 failures per one million missions and, of those, only 200 can be “uncontained failures” meaning that the engine comes apart and potentially threatens other vehicle elements.  So, all over the engine system we’re pushing more propellants, operating at higher pressures, generating more thrust, and squeezing out more performance efficiency, and we have to do this in a manner that results in an engine that has over twice as reliable as the heritage design.  The result is an engine that is bigger and heavier than its historical antecedent:

So, in summary, here are the components that we had to change to meet J-2X requirements:
• Turbomachinery
• Main injector
• Main combustion chamber
• Nozzle
• Gas generator
• Added a nozzle extension
• Swapped the sequencer with a controller

What does that leave?  Valves?  Nope.  Because of the higher flowrates and pressures, we had to drop the heritage designs for the valves and go to a design more akin to the Space Shuttle Main Engine.  Ducts?  Nope.  Once you’ve changed all of these other things, you end up rearranging the connecting plumbing just as a matter of course.  Even the flexible inlet ducts were changed slightly to accommodate more stringent design standards. 

Form follows function; function flows from requirements; requirements flow from mission objectives.  Different mission, different requirements, different function, and a different result.  Thus, the J-2 and the J-2X share a name and share a heritage — in many ways the J-2 (and the J-2S) was the point of departure for the J-2X design — but the J-2X is truly its own engine.  Lesson learned: Don’t assume too much from a name.


Inside The J-2X Doghouse: Beyond the Gas Generator Cycle

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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.