Tag Archives: Space Shuttle Main Engine

LEO Progress: RS-25 Adaptation

Posted on by .

I love dictionaries (yes, I know, you’re shocked; shocked!).  I have several at home and at work including a two-volume abridged Oxford English Dictionary (OED) that was a wonderful gift from my mother several years ago. 


The definitions and word origin above comes from the OED on-line site.  Embedded within our language is so much condensed history and accumulated knowledge that it’s amazing.  While I have no doubt that this is true of every language, I only know my own to any significant degree.  Indeed, I’ve been babbling my own language for a long time now — forty-some years — but I can always pick up a dictionary and learn something new with just the flick of a page or two.  You really can’t say that about too many other things.

As the title for the article and the definitions above suggest, for this article we’re going to talk about “adaptation,” specifically about adaptation of the RS-25 engine.  As part of the Space Launch System Program, we are undertaking something a bit unusual for the world of rocket engines.  We are taking engines designed for one vehicle and finding a way to use them on another vehicle.  Now, this is not a completely unique circumstance.  The Soviets/Russians really were/are masters of this kind of thing.  But for us, it is not something that we do very often.  In terms of the big NASA program rocket engines, I can only think of the RL10 that was originally part of the Saturn I vehicle as an engine design that has had a long second and even third career on other vehicle systems.  The truth is that engines and vehicles are, for us, generally a matched set and the reason is that we just don’t frequently enough build that many of either.  Note that, technically speaking, we are also adapting the J-2X from a previous program, Constellation/Ares, but that’s obviously a bit different in scope and scale given where we are in the development cycle.

The name “RS-25” is, as I’ve mentioned in past articles, the generic name for the engine that everyone has known for years as the “SSME,” i.e., the Space Shuttle Main Engine.  There is much to talk about the RS-25.  Lots and lots of stuff.  More stuff that I could possibly fit into a single article.  Here are just some of the RS-25 topics that we’ll have to defer to future articles:

       •  History and evolution
       •  A tour of the schematic
       •  Engine control, performance, and capabilities

For this article, I want to just talk about the scope of work that will be necessary to adapt RS-25 to suit the Space Launch System Program. 



Clear Communication
The most significant thing that has to be done to the RS-25 to make suitable for the new program is that it has to be able to respond to and talk to vehicle.  Remember, the Space Shuttle was developed in the 1970s and first flown in the 1980s.  Yes, many things were updated over the years, but given the lightning-fast speed of computer evolution and development, it is not surprising that what RS-25 is carrying around a controller basically can’t communicate with the system being developed today for the SLS vehicle.  The SLS vehicle would say, “Commence purge sequence three,” and the engine would respond, “Like, hey dude, no duh, take a chill pill” and then do nothing (my lame imitation of 1980s slang as best I remember it).


But here are the neato things that we’ll be able to do: We can use almost all of the work now completed on the J-2X engine controller hardware to inform the new RS-25 controller and we can use the exact same basic software algorithms from the SSME.  Because the RS-25 has a different control scheme from the J-2X, we cannot use the exact controller unit design from J-2X, but we can use a lot of what we’ve learned over the past few years.  And, because we can directly port over the basic control algorithms, we don’t have to re-validate these vital pieces from the ground up.  We just have to validate their operation within the new controller.  That’s a huge savings.

This is work that is happening right now, as I’m typing.  Pratt & Whitney Rocketdyne, the RS-25 developer and manufacturer, is working together with Honeywell International, the electronic controller developer, on this activity.  Within the next couple of months, they will have progressed beyond the point of the critical design review for both the hardware and software. 

In the long term, it is our hope that we will evolve to the utopian plain of having one universal engine controller, a “common engine controller,” that can be easily fitted to any engine, past, present, or future.  Such a vision has in mind a standardization of methods and architecture such that we could largely minimize controller development efforts in the future to the accommodation of obsolescence issues.  The simple truth is that development work is always expensive.  It would be nice to avoid as much as that cost as possible.  With the new RS-25 controller, we’re getting pretty close to that kind of situation.

Under Pressure
Have you ever spent much time thinking about water towers?  As in, for instance, why are they towers in the first place?  As you might have guessed, this is something that I wondered about many years ago as a pre-engineering spud.  They seemed to be an awfully silly thing to build when you could just turn on the faucet and have water spurt out whenever you want.  Ah, the wonderful simplicity of childhood logic:  Things work just because they do and every day is Saturday.

The reason that you go to the trouble of sticking water way up in a tower is so that you have a reliable source of water pressure that can absorb the varying demands on the overall system.  In the background, you can have a little pump going “chug, chug, chug” twenty-four hours per day pushing water all of the way up there, but by having this large, reserve quantity always in the tower, the system can respond with sufficient pressure when, at six in the morning, everyone in town happens to turn on the shower at the same time to start their day.  The pressure comes from the elevation of tower, the height of the column of water from top to bottom.

On the Space Launch System vehicle we have something of a water tower situation except that, in this case, we’re dealing with liquid oxygen.  In the picture below, you can see roughly scaled images of the Space Shuttle and the Block 2 Space Launch System vehicle.  On both of these vehicles, the liquid oxygen tank is above the liquid hydrogen tank.  This means that for the very tall SLS vehicle, the top of oxygen tank is approximately fifty feet higher relative to the inlet to the engines (as illustrated).  This additional fifty feet of elevation translates to more pressure at the bottom just as if it were a taller water tower.  However, in this case liquid oxygen is heavier than water (meaning more pressure) and the SLS vehicle will be flying, sometimes, at accelerations much higher than a water tower sitting still in the Earth’s gravitational field.  Greater acceleration also amplifies the pressure seen by the engines at the bottom of the rocket.


“So what?” you’re saying to yourself as you read this.  After all, higher pressure is a good thing, right?  If you have more pressure at the inlet to the engine, then you don’t need as much pumping power.  So, life should be easier for the engines with this longer, taller configuration.  There are two reasons why this is not quite the case.

First, think about when the vehicle is sitting on the pad at the point of engine start.  The pumps aren’t spinning so all the pressure you’re dealing with is coming from the propellant feed system.  And now, simply, for the SLS vehicle it will be different than before.  One of the tricky things about a staged-combustion engine, in general, is that the start sequence (i.e., the sequence of opening valves, igniting combustion, getting turbopumps spun up) is touchy.  Given that the RS-25 has two separate pre-burners — and therefore three separate combustion zones — and four separate turbopumps, the RS-25 start sequence especially touchy.  You have to maintain a very careful balance of combustion mixture ratios that allow things to light robustly, but not too hot, and a careful balance of pressures throughout the system so as to keep the flow headed in the right direction and keep the slow build to full power level as smooth as possible.  We have an RS-25 start sequence that works for Space Shuttle.  Now, for the SLS vehicle, we will have to modify it to adapt to these new conditions.

The second issue to overcome with regards to the longer vehicle configuration and the liquid oxygen inlet conditions is due total range of pressures that the engines have to accommodate.  When sitting on the launch pad, you have the pressure generated by the acceleration of gravity.  During flight, as you’re burning up and expelling propellants, the vehicle is getting lighter and lighter and you’re accelerating faster and faster, you can reach the equivalent of three or four times the acceleration of gravity.  So that’s the top end pressure. 

On the low end, you have the effect of when the boosters burn out and are ejected at approximately two minutes into flight.  When this happens, the acceleration of the vehicle usually becomes less than the acceleration of gravity meaning that the propellant pressure at the bottom of the column of liquid oxygen can get pretty low.  Momentarily, the vehicle seems to hang, almost seemingly falling, despite the fact that the RS-25 engines continue to fire.  Pretty quickly, however, the process of picking up acceleration begins again.  (In the movie Apollo 13, they illustrated a similar effect with the separation of the first and second stages of Saturn V.  The astronauts are pushed back in their seats by the acceleration until, boom, first stage shutdown and separation happens and they’re effectively thrown forward.  With the Space Shuttle and the SLS vehicle, however, the return to acceleration is not as abrupt as it was on Saturn V where they show the astronauts slammed back into their seats with the lighting of the second stage.)  The point is that the RS-25 has to accommodate a very wide range of inlet pressures while maintaining a set thrust level and engine mixture ratio.  While this has always been the case for the SSME/RS-25, the longer SLS vehicle configuration simply exacerbates the situation.

You’ll note that I’ve not talked about the liquid hydrogen here.  That’s because, as I’ve mentioned in the past, liquid hydrogen is very, very light.  Think of fat-free, artificial whipped cream.  Yes, the top of the hydrogen tank is much higher, but due to the lightness of the liquid, it doesn’t make much difference at the engine inlet even when the vehicle is accelerating at several times the acceleration of gravity.

Some Like it…Insulated
Look again at the pictorial comparison of the Space Shuttle and the Space Launch System vehicles shown above.  Do you see where the SSME/RS-25 engines are relative to the big boosters on the sides of the vehicles?  On the Space Shuttle, the engines were on the Orbiter and they were forward of the booster nozzles.  On the SLS vehicle, there is no Orbiter so the engines are right on the bottom of the tanks and their exit planes line up with the booster nozzle exit planes.  In short, the engines are now closer to those great big, loud, powerful, and HOT boosters.  We are in the process now of determining whether this poses any thermal environments issues for the RS-25.  Thus far, based upon analyses to date, there do not appear to be any thermal issues that cannot be obviated through the judicious use of insulation.

Other environments also have to be checked such as the dynamic loads transmitted to the engine through the vehicle or the acoustic loads or whatever else is different for this vehicle.  The point is not that all of these environments are necessarily worse than what they were on the Space Shuttle.  It is only that they all need to be checked to make sure that our previous certification of the engine is still valid for all of these considerations.

The Tropic of Exploration
So those are the three most obvious and primary pieces of the RS-25 adaptation puzzle: a new engine controller, dealing with different propellant inlet conditions, and understanding the new vehicle and mission environments.  Each of these pieces carries with it analysis and testing and the appropriate documentation so there is plenty of work scope to accomplish.  We are extremely lucky to be starting with an engine of such extraordinary pedigree, performance, and flexibility. 

Henry Miller once said, “Whatever there be of progress in life comes not through adaptation but through daring.”  It is our intent to prove Mr. Miller wrong in this case.  We will make progress by using the adaptation of RS-25 to enable the daring of our exploration mission.

Inside The J-2X Doghouse: Performance Measurement, Part 1 of 2

Posted on by .

I’ve had jobs of some flavor almost continuously since I was fourteen years old.  From delivering newspapers to cutting grass to flipping burgers to showing movies to mixing giant vats of coleslaw to instructing an aerodynamics laboratory course to hand counting vehicular traffic to researching the derivation of combustion stability equations, before I finally settled into something resembling a career (and while I was winding my way through the maze of secondary and higher education), I did all kinds stuff.  But when I did get my first job after graduate school it was doing test data reduction and performance calculations for the Space Shuttle Main Engine (SSME).  

How cool is that?  Trust me: very cool. I consider myself to be extraordinarily lucky in this regard.  This article is going discuss engine performance measurements and so it’s going to reach back to my very roots in this business. And that sounds like fun!

In past articles we’ve discussed rocket engine performance. We’ve talked about the “Big Three” operational points and performance measures that characterize an engine: thrust, mixture ratio, and specific impulse. Okay, but how do we know what these values are for any given engine?  I mean, we can do calculations with analytical equations, formulas, algorithms, or models that tell us what these parameters ought to be for a given rocket engine design; but when I’ve got an actual rocket engine sitting in front of me — a big, shiny, complex hunk of metal standing ten feet high and weighing thousands of pounds — how do I know how it actually functions and performs? 


Well, duh, you test it. Of course!  But making smoke and fire (and steam) does not, by itself, give you any data. The most that you could say from just watching an engine test is that it’s really, really noisy and that it makes a really, really big exhaust plume.  So, more than just observe, you have to take measurements during the test.  That’s how you get data.  As I’ve said many times, there are only two reasons to conduct rocket engine tests:  (1) to impress your friends, and (2) to collect data.  In order to get data on the “Big Three,” you need to measure thrust and you need to measure propellant flowrates. For this article, we’re going to focus on propellant flowrates. I will talk about thrust measurement in the next article.

Propellant flows are measured on a rocket engine test stand with “propellant flowmeters.”  Makes sense, right?  But calling something a “meter for flow” doesn’t tell you how it works. That’s like saying, “How do I make popcorn?”  Answer: “With a popcorn maker.”  No kidding. Thank you for playing and you’ve conveyed no useful or interesting information.


There are a number of different ways to measure fluid volumetric flow.  The units that we use for the very large flowrates feeding an engine are turbine flowmeters.  Have you ever blown into a small fan that’s turned off or perhaps a pinwheel?  If you blow hard enough, you can make the fan or pinwheel spin.  That is, quite simply, how a turbine flowmeter works: it’s a fan, i.e., turbine, stuck in a tube that spins as fluid flows through it. The faster the fluid flows, the faster the turbine spins. The thing that we measure is the speed of the turbine spinning.  The turbine has a number of blades (just like that small fan that you blow into).  We pick a spot on the tube in which the turbine sits and count how many blades pass by.  If we count, say, ten blade passes in a second, then there is more flow than if we’d only counted eight blade passes in a second. 



So, how do we count blade passes?  Well, there’s a little window in the side of the propellant duct and we sit a young college co-op in front of the window with a little hand clicker and scream “GO!” from the blockhouse… Okay, I’m fibbing.  We don’t treat our co-ops nearly that bad.  Usually.  Besides, there would be no way that the human eye and brain could keep up since we’re talking about hundreds of blade passes per second.  Instead, we measure it electronically.  Each blade contains a magnet in the tip.  The sensor on the outside of the tube is activated by the magnet.  Each magnet pass generates an electronic pulse or blip — what we call a “pip” — and we keep a continuous count of these accumulated pips.  The pip count is then recorded with each time step in the data collection process.  Then, after the test, we can translate this ever-increasing pip count into a pip-rate based upon these recorded times.  Mathematically speaking, the pip-rate at any given point is the slope of the pip-count plotted against time.


In order to translate a pip-rate into a volumetric flowrate, such as gallons per minute (gpm), the flowmeter needs to be calibrated.  We need to know how much flow is required to generate a blade passage, i.e., a pip.  If, for example, we knew that the passage of one gallon was enough to move the turbine exactly one blade pass of rotation, then a measured 100 pips-per-second would equal 100 gallons-per-second, or 6,000 gpm.  Thus, calibration of a flowmeter consists of flowing a known volume of fluid through the meter and counting the pips read: 


The truth is that it’s a bit more complicated in that the calibration varies with the speed of the turbine due to kinetic and mechanical issues of the rotating hardware and due to fluid dynamic effects of the fluid interacting with the turbine blades.  However, these are secondary effects as compared to the simple notion of figuring how much a pip is worth in terms of volume.

Luckily enough (or, really, strategically enough), the engine test stands are themselves set up to function as a calibration facility for the flowmeters.  This is because the propellant tanks have a known geometry and are equipped with fluid level sensors. 


As shown in the figure above, if we know at a particular time the height of the fluid in the tank and then, at a later time, we know a lower height of the fluid, then, using tank geometry, we know the volume of fluid that exited the tank and ran through the flowmeter.  In practice, we actually perform this calibration during an engine test.  That way we can be assured that the flowmeter rotor is spinning at a speed representative of where we’ll need measurements.

An observant reader would note here that if we know the volume consumed over time just from the level sensors in the tank, then we don’t need a flowmeter in the middle.  All you need is volume divided by time, right?  The problem is one of fidelity.  Because the level sensors are discrete points on the pole submerged in the tank, the measures of volume used for calibration are relatively big chunks, as in enough propellant to run the engine tens of seconds.  In order to get a decent calibration across several discrete level sensors, we typically need to run between 100 and 150 seconds of steady, mainstage engine conditions. The use of a calibrated flowmeter allows you to see variations in flowrate at much smaller time increments and this allows us to collect and observe more data with regarding to engine characterization at different conditions. You can almost think of the flowmeter as a useful interpolation tool between large chunks of time and consumed propellant.

You will note that so far we’ve just talked about volumetric flowrates and yet, when we talk about engine performance we refer to mass flowrates.  The difference between the volume and the mass of something is its density.  For our very pure propellants, fluid density is simply a function of fluid temperature and static pressure. So, we take temperature and pressure measurements immediately downstream of the flowmeter and, using either an interpolated look-up table or empirical curves, we can get density.  So, you put it all together and you end up with something along the lines of the following:


That is how you measure and calculate the mass flowrate of the propellants flowing through the feedlines and going into the engine using a turbine flowmeter.  The item from the “Big Three” to which this can be applied directly is the engine inlet mixture ratio, which is defined as the oxidizer mass flowrate divided by the fuel mass flowrate.

However, depending on the engine and vehicle design, not all of the propellants that go into an engine go overboard.  Often, warmed propellants are returned from the engine to the stage to act as pressurizing gases for the stage propellant tanks.  On the Space Shuttle, both gaseous oxygen and gaseous hydrogen were flowed back to the stage for this purpose. The rocket equation that essentially defines the parameter we know as specific impulse is only concerned with propellants that leave the vehicle so for specific impulse calculations you need to use inlet mass flow minus pressurization flow.  


As compared to the engine inlet mass flowrates, which for large rocket engines can amount to hundreds of pounds-mass per second, the pressurization flowrates are typically less than one or two pound per second.  Flows this small are more effectively measured using flowmeters different from the turbine flowmeters I’ve described above.  For our engine testing we use Venturi meters for these small flows. Venturi meters use a variable flow area coupled with pressure measurements to feed Bernoulli Equation relationships between pressure and fluid velocity. Once you know the fluid velocity, fluid density, and fluid flow area at any point, you can then calculate mass flowrate (for now, at least, I’ll not go any further with Venturi meter calculations).

This, then, wraps up the story with regards to propellant mass flow measurements and calculations on the engine test stands.  In the next article, we’ll go into the measurement of and calculation of thrust.  All of this discussion reminds me so much of my first days/weeks/months on the job working with SSME test data.  At first, it was just a bunch of bewildering numbers and data reduction tools and rules and calibration factors and work procedures.  I had no idea what was going on.  But gradually, as I dug into the data and talked to people and dissected the computer codes and tools we used, I began to piece it all together as to what these measurements and calculations actually meant. Seemingly every day brought a new epiphany in understanding.  Boy oh boy, that was fun!

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

Posted on by .

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.

Where:
      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.

Where:
      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. 

 

Inside The J-2X Doghouse: Engine Control — Open versus Closed Loop

Posted on by .

As I was driving to work this morning, I came up over a rise and saw suddenly appear in my windshield, over towards the left on the other side of the road, a police cruiser with a radar gun mounted in the window.  Even before I could think about it, the pressure of my right foot on the accelerator lessened.  I then instinctively looked at the speedometer and found that I was traveling 48 miles per hour on a road for which the speed limit is 45 mph.  Thankfully, the officer apparently forgave me the 3 mph violation and continued to wait where he was for a better opportunity to serve and protect the community. 

 

What I find interesting about that little episode was the immediate, unthinking response I made in response to seeing the cruiser.  In terms of control systems, that could be called a feedback loop.  My senses received data, my brain rapidly processed that data and then sent a signal to react to the data, and then my calf muscles responded by easing up on the throttle.  The car, in turn, slowed to respond to the lower throttle.  Never mind, of course, that if I’d been really, really speeding all this would have been too late (and I would have arrived at work very angry), there was nevertheless a closed-loop response that is not too dissimilar to what we do for rocket engine control … in some cases.

 

First, let’s get familiar with a couple of terms –

 

Open-loop: We typically refer to something as open-loop when we have instrumentation that measures conditions in the engine but the engine itself does not respond to those measurements.

 

Closed-loop: We typically refer to something as closed-loop when we have instrumentation that measures conditions in the engine and then, potentially, the engine takes action based upon those measurements.

 

It is based upon those definitions that I would call my response to seeing the cruiser closed-loop since I responded and did something with the data.  Another example would be the more modern systems that are used to monitor and control automobile systems.  It used to be that you had a temperature measurement stuck in the coolant loop.  You could watch the temperature rise, but until the system went kaput and boiled over leaving you stranded alongside the road, there was no active, closed-loop control.  Nowadays, if the computer in my pickup truck sees that the engine temperature is too high, it will take action to try and protect itself.  For example, it will inhibit the use of the air conditioning system since that represents an additional power requirement on the engine.



 So, what do we control on a rocket engine?

 

One thing that we control is power level.  On the Space Shuttle Main Engine (SSME), power level is controlled in a closed-loop manner.  This means that the main combustion chamber pressure is measured as an indication of thrust level and in response to that measurement a valve is opened or closed to increase or decrease the engine power level.  On the J-2X, power level is controlled in an open-loop manner.  This means that we measure the main combustion chamber pressure but we don’t have any feedback loop where we control a valve to ensure that we’re on target.  Instead, should we happen to be off on power level, we have to physically change an orifice in the engine between tests.  The “feedback loop” is data analysis and a guy with a wrench.  Which approach you choose to take are dependent upon your requirements of performance and affordability. 

 

Another thing that we control on a rocket engine is the mixture ratio (i.e., the ratio of oxidizer to fuel).  Given that on a rocket you are carrying both your oxidizer and fuel with you in the vehicle, you certainly want to make sure that you consume your propellants in the correct ratio to get the most uumph out of them.  Again, on SSME we control mixture ratio in a closed loop manner.  There is actually a small flowmeter on the SSME and, using the data from that flowmeter (and some associated calculations), we move a valve on the engine to dial in the correct mixture ratio.  It’s a pretty nifty system.  Also again, on the J-2X, we have an open-loop system for mixture ratio just like we have for power level.  We test the engine, look at the results, and, if necessary, make a physical change to the engine in the form of an orifice.

 

Because of these two areas, power level and mixture ratio, SSME is usually referred to as a “closed-loop engine” and J-2X is usually referred to as an “open-loop engine.”  Now, this terminology is not entirely correct since there are some closed feedback loops within the J-2X control system pertaining to engine health and status diagnostics, but we all know how enduring shorthand designations can be.  Also, engines don’t have to be one or the other.  They can be half-and-half.  The engine used on the Delta IV vehicle, the RS-68, sort of falls in this category. 

 

How you choose to design your engine control system is driven by your requirements.  Put real simply:  The SSME is all fancy-schmancy because it had extremely tight power level and mixture ratio precision requirements and because it was a reusable engine.  The J-2X is intentionally more simplistic because it has looser precision requirements and because it is expendable (and throwing away orifices is a whole lot cheaper than throwing away valves if your requirements will let you get away with it).  Requirements drive design.

 

Note that I will save the fun topic of engine diagnostics — and the potential for long philosophical meanderings within that realm — for future posting. 


 


 

 

Let’s end this posting with a fun little exercise.  Above is a simplified schematic of a gas-generator cycle engine kind of like a J-2X.  I have shown in the schematic two orifices #1 and #2 (highlighted in yellow).  With those two orifices, we can calibrate the engine.

 

Scenario:  Power level too low, i.e., measured main combustion chamber pressure too low.

·         Solution:  Increase the size of orifice #1.

·         Explanation:  By increasing the size of orifice #1, I will deliver more oxidizer to the gas generator.  This will deliver more power to both turbines thereby increasing how much propellant gets pumped into the engine.  More propellants pumped in equals more thrust and greater overall power level.

 

Scenario: Power level too high, i.e., measured main combustion chamber pressure too high.

·         Solution:  Decrease the size of orifice #1.

·         Explanation:  The exact opposite of the previous scenario.

 

Scenario:  Mixture ratio too low, i.e., the flow of oxidizer is too low in proportion to the flow of fuel, as measured by the test facility.

·         Solution:  Decrease the size of orifice #2 and decrease size of orifice #1.

·         Explanation:  By decreasing the size of orifice #2, I decrease the amount of flow that is diverted around the oxidizer turbopump turbine.  I therefore increase the flow through the turbine thereby increasing pumping power of the oxidizer side.  So I increase oxidizer flow to the engine.  However, by increasing oxidizer flow to the engine and doing nothing else, I’ve probably messed up my overall engine power level so I’ve got to back down a little bit by decreasing the size of orifice #1.

 

Scenario:  Mixture ratio too high, i.e., the flow of oxidizer is too high in proportion to the flow of fuel, as measured by the test facility.

·         Solution:  Increase the size of orifice #2 and increase size of orifice #1.

·         Explanation:  The opposite of the rationale for the scenario immediately above.

 

See, being a rocket scientist isn’t that difficult, really.  Now you too can calibrate an open-loop rocket engine.

 

 

P.S., I read in the paper this morning that NASCAR racer Kyle Busch had his civilian driver’s license revoked for 45 days for doing 128 mph in a 45 mph zone.  Well, at least I wasn’t going that fast when I came over the rise this morning and saw the police cruiser.  Then again, I wasn’t driving a $400,000 Lexus LFA sports car like Mr. Busch was…

 


 

 

 

J-2X Progress: Valves, Commands into Action

Posted on by .

Everyone seems to like analogies between the composition of a rocket engine and that of the human body.  These are often colorful but not always helpful.  In some cases, however, they work pretty well.

Okay, so let’s start with your body as it is.  Now, imagine removing all of your bones.  Guess what?  You’re an immobile lump.  Even if your brain is sending signals and your muscles are contracting, you’re not really moving anywhere.

This time, let’s instead start with your body as it is, but now imagine removing all of the muscles and tendons that connect the muscles to the bone.  You’ve got a central nervous system and you’ve got bones, but with nothing to flex, the chain is broken and you’re stuck where you sit (assuming that you can still actually sit).

And, of course, if you instead start with your whole self and imagine removing your brain and/or your central nervous system that connects your brain to your muscles, again, you’ve achieved perfect immobility (i.e., you look like me on Saturday afternoons during college football season).

The point is that in order for you to be up and about, shoveling snow, doing laundry, playing pool, typing, whatever, you need both the command center that figures out what signals to send — your brain — and you need things that turn those signals into action — your muscles and tendons and bones.  In a rocket engine, the analogue for the brain is the engine controller.  It is a computer that receives instructions from the vehicle and sends out commands to the engine pieces so as to fulfill those instructions.  The analogue for the muscles are the valve actuation systems.  These are the things that “flex” and cause movement.  And the analogue for the bones, the final effectors that make things happen, are the valves.

The controller sends out signals and then the actuation system responds by shuttling pressurized working fluid — helium for J-2X though some engines use hydraulic fluid instead — where it needs to go so that the valves move and the engine comes to life.  The engine goes from being a lump of inert, shiny metal to a “living” beast of flowing propellants, spinning turbomachinery, lots of fire, and thundering, rumbling thrust.

On the J-2X, there are 42 valves.  Most of this number is made up of small valves like check valves, solenoid valves, and valves in small lines like the bleed lines.  There are also a handful of big valves — the primary valves — that directly control the flow of propellant and, in one case, combustion products along the plumbing of the engine.  Each of these primary valves is connected to a valve actuator, i.e., the muscle.  These valve actuators convert the energy of high pressure helium gas into mechanical rotation of the valve.  This is accomplished by pressurizing cavities and moving pistons and, in this way, the valve is pushed opened or closed.  I’ve used this schematic shown below before, but it is useful here as well since it illustrates the primary J-2X valves: Main Fuel Valve (MFV), Main Oxidizer Valve (MOV), Gas Generator Fuel Valve (GGFV), Gas Generator Oxidizer Valve (GGOV), and the Oxidizer Turbine Bypass Valve (OTBV).

The control logic for J-2X is relatively simple.  The whole subject of different kinds of control logic is a good topic for a future article, but suffice it to say that for normal operation the J-2X: starts on command, can change between two power levels on command, and shuts down on command.  The control system is designed to do other things as well, including monitoring the health of the engine, but these operations are the commanded functions.  Start and shutdown can be simplistically thought of as: the valves open and the valves close.  It’s a bit more complicated since the timing of opening and closing is extremely important, but the open/close notion is basically true.  The oddball action is the one consisting of changing power levels.  That is accomplished by controlling the power to the oxidizer turbine via the OTBV.  This bypass valve effectively allows for limited, independent control of the two turbopumps.  By altering the power to the oxidizer turbopump (OTP), you can control the engine thrust level (and, simultaneously, mixture ratio).

The OTBV for J-2X is designed and built by Pratt & Whitney Rocketdyne (PWR), the prime contractor for the whole engine.  In addition to being responsible for the “oddball action” on the engine of changing power levels, it represents a challenging design due to the range of operating conditions.  Unlike the other primary valves on the engine that see, essentially, one narrow range of environmental conditions, the OTBV has to function in temperatures approaching 420 degrees below zero Fahrenheit (liquid hydrogen conditions) immediately prior to start and then, suddenly, within 1 second of ignition of the gas generator, see temperatures approaching 750 degrees above zero Fahrenheit (combustion products).  That broad range of operating conditions requires special design considerations and special materials.  Not only do you have to worry about wear and tear under such harsh conditions, but you also have to think about simple operation under the extremes of thermal expansion.

The original, Apollo-era J-2 engine also had an OTBV, but it was used slightly differently and was designed much differently.  It was a butterfly valve whereas the J-2X OTBV is a ball valve. 

No, the valves shown in the picture are NOT rocket engine valves.  I can’t show any internal workings of rocket engine valves.  In fact, I am not even allowed to describe the general design details that make the J-2X OTBV kind of unique.  However, the basic elements of rocket engine valve functionality for butterfly and ball valves are essentially the same as these water valves.  The biggest difference is the replacement of the handles with pneumatically driven actuators.  Back during the Apollo era it would seem that butterfly valves were most frequently used, but after many years of usage on the Space Shuttle Main Engine, ball valves are often preferred these days.  They generally require less torque to move and they generate better flow characteristics and flow rate control capability.

The first OTBV unit for use on the upcoming development engine testing for J-2X is in the later phases of manufacturing at the PWR in Los Angeles.  All of the individual piece parts are schedule to be complete by the beginning of February and assembly will begin the middle of February.  The valve then will be integrated the actuator and shipped to the NASA Stennis Space Center to be put on the first engine.

J-2X Extra: The Faces Behind J-2X, NASA MSFC, Part 1

Posted on by .


Years ago, I worked in support of the Space Shuttle Program, specifically on the Space Shuttle Main Engine (SSME) from the engineering side of the house.  I was an analyst and sometimes Datadog.  By happenstance, I learned from another relative that whenever a launch occurred my mother, back home in Pennsylvania, would routinely tell folks at work or at church or wherever that her son was responsible for that Shuttle launch.  I said, “Mom, don’t tell people that.  I’m just one of hundreds or, really, thousands of people behind the whole thing.”  She replied, with biased motherly wisdom, “Yes, but you’re all responsible for doing a good job so that the whole thing works, right?”

Lesson learned #1:  Don’t argue with Mom.

Lesson learned #2:  She was right.  Any venture as big and complex as Shuttle, or even as big and complex as the development effort for J-2X really does require that everyone does their job well.  In that way, we’re all responsible.

So, I’m going to introduce you now to just a few of the people (out of hundreds) responsible for making J-2X a reality.  Right now I’ll focus on the top leaders Upper Stage Engine Element office here at NASA Marshall Space Flight Center.  In another posting, I’ll tell you all about the subsystem managers.  And perhaps, later if I’m lucky, I’ll sneak some pictures of the good people out at Pratt & Whitney Rocketdyne as well.

So, here we go…

First, we have our Element Manager, Mike Kynard.  He is a graduate of the University of Alabama [Roll Tide!], grew up not far from Tuscaloosa, and, to put it mildly, is a devoted fan of Crimson Tide football.  He started working at NASA MSFC as a co-op in college back and accepted a full-time job in 1985.  He spent a good spell supporting SSME from the resident office at NASA Stennis Space Center in Mississippi in the 1990s and eventually rose to become the deputy project manager for SSME back in 2004.  He has two beautiful little girls and, in addition to everything else, somehow finds the time to play on a team with me in a billiards league.  Sometimes, on rare occasions such as a blue moon, he even manages to beat me in a game of 8-ball.

Mike’s deputy is Tom Byrd.  He is a graduate of Memphis State University who, once upon a time long, long ago was a competitive bicyclist.  He started working here at NASA MSFC in 1983 in the area of valves and actuators, specifically supporting SSME (yes, you’ll see SSME as a recurring theme both here and when later tell you about the subsystem managers).  Along the way since then, he was a subsystem manager for the Fastrac engine development effort (a NASA MSFC in-house project), was the NASA chief engineer for COBRA engine development effort, and he supported the Shuttle program in the area of systems engineering and integration.  Plus, back in 1994, representing the NASA engineering community, he spent a total of four weeks in Russia studying their space transportation technologies.  Tom has one young son who, given the uncanny resemblance, we suspect might be the product of a direct cloning experiment … but we don’t talk about that too much.

Our Chief Engineer is Eric Tepool.  His job is to function as the balance point between technical and programmatic aspects of engine development.  He also indirectly functions as a balance in another way, relative to our element manager’s tendencies, in that he is a graduate of Auburn University [War Eagle!].  After making his mark as a star high school athlete, just as his father was a star high school and collegiate athlete, Eric turned down athletic scholarship offers to settle into the pursuit of engineering.  He started at NASA in 1990 in the turbomachinery branch, supported development and certification of the two new turbopumps for the Block 2 SSME configuration, moved on to be a subsystem manager for the COBRA engine development effort, resident manager for the Fastrac project, and the NASA-side lead systems engineer for the Integrated Powerhead Demonstrator project.  He has two kids currently in college (one at Alabama, one at Auburn … yikes!) and one soon to be on her way to college.

The other “technical balance point,” in accordance with the NASA governance model, is our Chief Safety Officer Phil Boswell.  He’s been working here at NASA MSFC since 1985.  He started right after graduating from the University of Alabama at Birmingham (UAB).  While he started in the Safety and Mission Assurance Directorate, and he’s back there now, in the interim he spent many years working within the Engineering Directorate on such projects as microgravity experiments for Shuttle and the MIR space station, the Orbital Space Plane program, and just about every propulsion element of the Shuttle itself including engines, tanks, and boosters.  Phil played on the college tennis team, plays tennis still, and is an excellent golfer.  His son, who will be attending UAB starting in the autumn (pre-med) follows in his father’s footsteps and is also an excellent golfer.  Little known fact about Phil: he loves swing dancing and even took a year of lessons with his wife.



So, that’s the top leadership group for the J-2X development effort.  Good guys, hardworking engineers and managers, all around.

 

Inside The J-2X Doghouse: The Gas-Generator Cycle Engine

Posted on by .


Welcome back to the J-2X Doghouse.  The last time that we met here, we discussed the fundamentals of what exactly makes something a rocket.  As I explained, on the conceptual level, rockets aren’t really “rocket science.”  You get the propellants together, light them on fire, and eject them out the back end of the vehicle.  Simple enough.
 
Okay, but how do you move that much propellant and make that much smoke and fire, enough to propel something as big as, say, the Saturn V that was over 300 feet tall and weighed millions of pounds?  That’s where things get interesting and technically difficult.  As I said before it is all a matter of power. And to get power you use an engine.

What makes a rocket engine an engine is the fact that it contains more than just a combustion chamber where the propellants mix.  It is an arrangement of machinery that, once started, feeds and powers itself.  During operation, a rocket engine uses some cycle – some circuit of piping and thermodynamics and combustion and valves and control system and rotating machinery – to keep itself up and running and generating thrust.

Think about your car engine.  You turn the key, the engine gets up and going, and then it can sit there for hours idling, running happily all by itself, converting gasoline and air into mechanical energy, with no additional input from you.  You don’t have to manually pump the gas into the injectors (or the carburetor).  You don’t have to plug it into an outlet to feed it more electrical energy.  It’s self-sufficient until you turn it off or until you run out of gas.  That’s what truly makes it an engine.  It’s similar with a rocket engine except that the product is not mechanical energy; the product is very fast moving gases generating lots of thrust.

 

For rocket engine conceptual design, in terms of making it an engine, the goal is always, “How do you keep the pumps pumping?”  These are extremely powerful pumps moving lots and lots of fluid, so you need some powerful energy source to drive them.  The answer is to use what you’ve already got in the engine: the propellants.  There are different ways to do this and thus you have different engine “cycles,” i.e., component arrangements.  The most common rocket engine cycles are the gas-generator cycle (examples include J-2X, J-2, F-1, RS-68, and Vulcain 2 – see pictures above), the expander cycle (examples include RL10 and Vinci), and the staged-combustion cycle (examples include Space Shuttle Main Engine and RS-170/180).  In addition to these, there are many other cycles and variations as well.  Each different cycle has advantages and disadvantages and, usually, constraints linked to physics.  Choosing the right cycle to fit the mission application is generally the first decision that an engine designer has to make.  Because this is a blog dedicated to J-2X, I will focus on the gas-generator cycle engine.

Ideally, what you would want to do with a rocket engine is use all of your propellants in as efficient manner as possible meaning that you would want to use all them in the production of thrust.  In a gas-generator engine, however, you concede right up front to a loss of some efficiency to achieve greater engine simplicity.  You use a certain amount of the propellants brought into the engine almost entirely to keep the engine running rather than for generating thrust.  In practice what this means is that you have a separate, small combustion chamber within the engine that does nothing but produce gases to drive the turbines connected to the propellant pumps.  As compared to the large quantities of propellants being pumped through the whole engine, the amount going to the gas generator is small (less than 3% for J-2X), but once used to drive the turbomachinery, the exhaust is drained of much of its thrust-generating energy. 

Below is a simplified schematic of a gas-generator cycle rocket engine like the J-2X.  The propellants, liquid hydrogen (fuel) and liquid oxygen (oxidizer), enter the engine and go immediately into the pumps: the fuel turbopump (FTP) and the oxidizer turbopump (OTP).  There, the mechanical energy of the spinning pumps is turned into high pressures in the liquid propellants. 

 

After exiting the pumps, a small amount of each propellant is tapped off to supply the gas generator (GG).  The GG is, in essence, a small rocket engine embedded within the larger rocket engine.  It makes hot, high-pressure combustion products, steam and gaseous hydrogen, that are used to drive first the turbine connected to the fuel pump and then the turbine connected to the oxidizer pump.  After driving the two turbines, this still-warm gas is used first to warm the helium flowing through the heat exchanger (HEX) that is used to pressurize the oxygen tank of the stage and is then dumped along the walls of the nozzle extension to keep that relatively cool.  The video below is a component test of the J-2X GG performed at NASA MSFC.  Even with the relatively small amount of propellant that the GG burns, an enormous amount of energy is released to drive the turbopumps.

The rest of the liquid oxygen coming out of the oxidizer pump, meaning that which is not going to the GG, is directed through the main injector and into the main combustion chamber (MCC).  The main injector is analogous to a fuel injector in a car engine except that here it injects two propellants through hundreds of injector elements.  The effectiveness of this injection and the mixing of the propellants are crucial for overall engine performance.

The hydrogen circuit after the fuel pump is more complicated.  This is because the hydrogen is used to cool the nozzle and combustion chamber walls.  The walls of these two components are essentially hollow.  They contain hundreds of passages for the hydrogen to flow thereby keeping the walls from melting due to the extreme high temperatures of the contained combustion zone.  After doing its job as coolant, the hydrogen is then directed through the main injector and into the MCC.  Not shown on the diagram is the fact that a very small amount of the warm hydrogen gas is tapped off prior to entering the main injector and is routed back to the stage to pressurize the hydrogen tank (like the helium through the HEX on the oxygen side).

It is in the MCC where the mixed hydrogen and oxygen combust to make steam and residual hydrogen gas.  The temperature of that combustion is approximately 6,000 degrees Fahrenheit and in the J-2X the pressure is approximately 1,300 pounds per square inch.  These combustion products are then accelerated to sonic velocity at the converging throat of the MCC and then to supersonic velocities down the diverging nozzle and nozzle extension.  As discussed previously, it is the high-velocity expulsion of these hot gases that produces thrust.

Note that the turbine exhaust gases dumped along the nozzle extension still generate some thrust, but not as effectively as the combustion products that are accelerated through the nozzle throat.  This loss of effectiveness is the price that you pay for this relatively simple engine cycle.  As a comparison to a more complex engine cycle, do a web search for the schematic for Space Shuttle Main Engine (SSME).

FYI, the other items denoted on that GG-cycle schematic above are the control valves: the main fuel valve (MFV), the main oxidizer valve (MOV), the gas-generator fuel valve (GGFV), and the gas-generator oxidizer valve (GGOV).  These primary valves, along with several other minor ones, are used to control the engine during the start and shutdown of the engine. 

So that’s how a gas-generator cycle engine like the J-2X works.  As this blog continues and as we head towards testing next year, I will continue to report on the progress of the components that make up the engine.

Inside the J-2X Doghouse: What is a Rocket?

Posted on by .


For as long as anyone can remember here at NASA’s Marshall Space Flight Center, the collection of engineers who analyze and evaluate rocket engine test and flight data results have been called “Datadogs.”  However, that time-honored moniker is a title that must be earned.  It’s not automatic based upon your job assignment.  It is based upon your ability to create a coherent technical narrative derived from hundreds of pieces of data spanning pre-start purge schedules, through engine start to mainstage operation, through shutdown transients and, finally, post-test inspections.  With every engine firing we ask: What happened and, more importantly, why?  The Datadogs provide the answers.

So, as a regular part of the J-2X Blog, I will be inviting you into the J-2X Doghouse just to ramble a bit about rockets and rocket engines in preparation for the upcoming J-2X development testing next year. 

The most basic question is, of course, what is a rocket?  Often, when lost in the mountain of ten thousand details of fabrication processes and assembly procedures and structural analyses and operational manuals and information of all flavors, even rocket scientists sometimes lose sight of the most basic concepts.  Yet any child who has ever blown up a balloon and then let it fly across the room as it deflates has experimented in rocketry.  A rocket is simply a vehicle that is self-contained and self-propelled.  It takes in nothing from its external environment and it achieves motion from Newton’s principle of a reaction resulting from every action.  A rocket effectively throws stuff out the back end while what remains in the rocket moves forward thereby balancing the net sum of inertia.

 

In technical terms, the balloon flying across the room — likely landing in your uncle’s soup thereby causing a minor family crisis — is a pressure-fed, mono-propellant rocket.  The stretchy plastic of the balloon supplies the pressure and the single propellant is the breath with which the balloon was filled.  The pressure from the plastic pushes the air out the back end.  The air goes one way rapidly and the balloon itself goes hurtling through space in the opposite direction.  Ta-da, a rocket!  And now you are privy to the NASA secret that rockets, at their most basic, conceptual level, are pretty darn simple.

So, what makes a rocket engine different than a child’s balloon?  Power.  In order to throw thousands of pounds of a launch vehicle into the sky and accelerate it to thousands of miles per hour, you need lots and lots of power.  Rather than relying on pressure to push the working fluid out the back end, a large rocket engine like J-2X uses very powerful pumps.  And, rather than relying on just the velocity generated by moving the fluids, a large rocket engine taps into the chemical energy released by combustion. 

For example, during every second of operation the J-2X pumps hundreds of pounds of hydrogen and oxygen into a chamber not much bigger than a large spaghetti pot.  There, these fluids combust, making steam (and residual hydrogen gas) at blistering hot temperatures of thousands of degrees.  That tremendous amount of energy is then directed out the back end, accelerating the hot gases down the length of the nozzle to supersonic speeds, converting thermal energy to kinetic energy all along the way. 

How much steam does this make?  Well, if you ever have the opportunity to see a J-2X engine test, bring an umbrella.  A full duration test will make enough steam to make its own rain cloud in the sky.  Below is a video of a Space Shuttle Main Engine test in stand A2 at NASA’s Stennis Space Center in Mississippi.  Tests of the J-2X will look quite similar.

 

Thus, the tough part about rocket engines is not their basic concept.  That’s simple.  The tough part is building a device that can harness the power necessary to make that simple concept useful.  As we go along, we’ll discuss that tough part in more detail.