J-2X Progress: Shaking up the Night

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The first rocket engine test that I ever saw in person at the NASA Stennis Space Center in southern Mississippi occurred well over twenty years ago.  I’d already been doing test data analysis and power balance analysis in support of the Space Shuttle Main Engine (SSME) Project for some months.  I had made several data review presentations in front of then chief engineer (and rocket engine legend) Otto Goetz.  I could quote engine facts, statistics, and tell you all about how the SSME worked.  I’d even seen some videos of engine tests.  But it was not until I saw a test in person that I achieved the state that can only be described as awestruck.

WVBIt was late in the evening and a little chilly.  Though we’d arrived at the control center before dusk, test preparations had dragged on so that now darkness had enveloped the center.  The test stand stood out against the blackened sky like a battleship docked in the distance.  Brian, my team lead at the time with whom I’d driven down from Huntsville, and I were standing outside in the control room in the parking lot.  The radiation from the hydrogen flair stack off to our left warmed one side of our face as the breeze cooled the opposite cheek.  The wailing of the final warning sirens drifted off and all that could be heard was the burning torch of the flair stack and, in the distance, the low surging and gushing of water being pumped into the flame bucket.  We were a just a couple of hundred yards from the stand.

Then, the engine started.

Picture1First, there is the flash and then, quickly, the wave of noise swallows you where you stand.  Unless you are there, you cannot appreciate the volume of the sound.  It is not mechanical exactly.  It is certainly not musical.  It is not a howl or a screech.  It is, rather, a rumble through your chest and a shattering roar and rattle through your head.  You think instinctively to yourself that something this primal, this terrible must be tearing the night asunder; it surely must be destructive, like a savage crack of thunder that continues on and on without yielding.  You are deafened to everything else, deprived of hearing because of all that you hear.  Yet before your eyes there is the small yet piercing brightness of the engine nozzle exit that can just be seen on what you know to be deck 5 of the stand and, to the right, there are flashes of orange flame stabbing into the billowing exhaust clouds mounting to ten stories high, tinged rusty in the fluctuating shadows.  It is like a bomb exploding continuously for eight minutes and yet the amazing thing, in incongruent fact so difficult to grasp as you are trying to absorb and appreciate the sensation is that the whole event is controlled and contained.  You cannot believe that so much raw power can be expressed by what is only a distant dot within your field of vision.

This is an experience that I wish everyone could have.  There are so many extraordinary feats of engineering all around us that we can appreciate and admire, but nothing for me has ever been as visceral as seeing an engine test, especially at night, with the nozzle open to the night air (and not buried in a diffuser).  No engine schematic or listing of characteristics or series of still pictures is an adequate substitute for the majesty of that controlled power.

Since that first test, I’ve seen any number of other engine tests including SSME (what we now call RS-25), a couple of other, smaller engines, and, of course, J-2X.  But it was not until the end of June of this year that I again had the opportunity to see a nighttime test on NASA SSC test stand A1.  This was J-2X E10002.  Below is the video, and it’s really cool, but I wish that you could have been there, standing beside me in the parking lot.  Listen carefully to the end of the recording and you’ll hear people cheering.  I was amongst the appreciative, awestruck chorus.

 

 

 

LEO Progress: J-2X to Test Stand A1

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“Nothing behind me, everything ahead of me, as is ever so on the road.”
– Jack Kerouac, On the Road

Recently, J-2X development engine 10002 was on the road.  If you remember, E10002 went through a six-test series on test stand A2 that began in February and finished up in April.  The next planned phase of E10002 testing is on test stand A1.  In between these series, the engine was back in the assembly area of NASA Stennis Space Center Building 9101.

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This respite between test series allowed for a complete series of inspections of the engine hardware.  This is vital piece of the learning process for engine development.  The basic truth is that a rocket engine is just darn tough on itself when it fires.  The reason that we test and test and test is to make sure that our design can stand up to the recurring brutal conditions.  The chance to look for the effects of that testing through detailed inspections away from the test stand is an opportunity to collect a great deal of useful information.

Also, even before the engine arrived at the assembly area, the stub nozzle extension was removed.  This was done while the engine was still installed in the test stand.  Remember, the testing on test stand A2 was performed with a passive diffuser and so we were able to use an instrumented stub nozzle extension to examine the nozzle thermal environments.  On A1, there will be no diffuser.  We’ll be firing straight into the ambient Mississippi afternoon and so we’ll not have any nozzle extension attached.  The other change made to the engine — this one made while in the assembly area — was that we swapped out the flexible inlet ducts so that we can use our specially instrumented ducts for the gimbal testing on A1.  These ducts will provide a great deal of unique data when we gimbal the engine and force the inlet ducts to twist and bend and they are designed to do.

Below is my favorite picture from the recent assembly activities for E10002 back in Building 9101.

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“FOE” stands for “foreign object elimination.”  I love this picture because it is a demonstration of how dedicated and meticulous are our assembly techs and how much basic integrity they bring to their job every single day.  In the rocket engine business we tend to be fanatics about foreign objects (i.e., random debris of unknown origin) hanging around.  The reason for this is that if you spend enough time in the business, you will eventually have a story of what happens when trash gets into the engine.  The rocket in question might be an amazingly powerful beast pumping out five hundred pounds of propellant per second generating 300,000 pounds of thrust, but all it takes is one bit of junk in the wrong place to destroy the whole thing in fractions of a second.

In the picture above is a small nut that was found in the periphery of the assembly area.  Shoot, if this was my garage you’d be lucky to find a clean patch not strewn with various bits of stuff like nuts, bolts, wads of duct tape, old hunks of sandpaper, that lost pair of pliers, a “Huey Lewis and the News” cassette, or, well, who knows what.  But the rocket engine assembly area is NOT my garage (thank goodness).  When something is found like a stray nut in the picture above, that sets off an investigation.  Where did it come from?  How did it get loose?  What procedure allowed this nut to escape control?  This is serious stuff.  Yet, just think about how easy it would have been for a tech to see that stray nut, pick it up, and stick it in his pocket.  They could have avoided the whole minor investigation thing entirely.  But that’s not what they did and that is not what they do.  Because they know that if they do not show the necessary integrity and methodical approach to continue to learn and perfect our procedures, then the next stray nut could be lodged where it could do terrible damage.

Here are the techs moving E10002 out of its assembly bay and unpacking it for transport.

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And a couple more pictures of the process in Building 9101.

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Here’s the engine ready for the road, then being lifted up the side of the test stand, and then sitting in the porch area while sitting on the engine vertical installer.  I really enjoy the pictures of the engine trucking about sunny NASA SSC.  That picture was the inspiration for including the Jack Kerouac quote at the beginning of the article.  It’s bright and shiny and full of so much thrill and promise.

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All of this should look reasonably familiar.  It is the process that we follow, more or less, whenever we take an engine out to the stand.  Getting the engine into and out of A1 is a bit easier since you don’t have to deal with moving the diffuser out of the way, but they’re really quite similar.  The slightly different flavor that we’ve got for this testing is the addition of the thrust vector control system.  In the picture below you can see where these hang.  The engine mounts up with the gimbal bearing to the stout, yellow thrust take-out structure.

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The two hydraulic actuators are also attached to the thrust take-out structure but slightly outboard and at 90 degrees apart.  These actuators are what will swing the engine around as if we were steering a vehicle.  Here, below, is another, closer view of the thrust take-out structure and the actuators.

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In the picture below, E10002 is mounted up to the thrust take-out structure.  The gimbal bearing is the shiny object towards which the arrow is pointing.  If you look over to the right side of the picture, you can see one of the “scissors ducts,” i.e., the flexible propellant inlet ducts.

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The next picture shows one of the hydraulic actuators hooked up to the engine.  As you can see the tolerances are awfully tight.  That’s an important vehicle integration consideration.  If this was a vehicle stage rather than a test stand to which we were attaching the engine, the thrust structure and the actuators would be the responsibility of the stage manufacturer.  Making sure that the stage and the engine can work together in such close quarters takes a great deal of vigilance between the two teams.

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So, you’re probably asking yourself, to what do these actuators connect on the engine?  That’s a very good question.  It certainly isn’t obvious from the assembly pictures.  The actuators connect to the forward manifold of the main combustion chamber (MCC).  Below is a computer model of the MCC with the two actuator attachment points shown.

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The MCC is really the heart of the whole engine, the sturdy framework stuck right in the middle, so it makes sense that when you want to push the engine around, this is what you’d have to push.  This final picture below of the engine completely mounted into the stand.  Again, it’s amazing to think about that whole thing being able to move about and not having one component run into another component or the actuators or the stand.  It is quite the integration miracle.

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Testing for E10002 on test stand A1 will commence in June.  So, if all goes well, for the next J-2X update, I’ll be able to link in a video of the engine firing and twisting about.

BTW, NASA is in the process of swapping software used for posting blog articles and comments and such.  As part of this process, they have to shut down the capability to accept input comments for a short time, specifically the first two weeks of July 2013.  Sorry about that.  But after that, it ought to be up and running as normal.

LEO Progress: RS-25 Adaptation

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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.

Liquid Engines Extra — Introducing LEO

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The following picture is a test.  Find me amidst the mess…

That’s right.  I’m the handsome chap with the snazzy specs.  See, I’m just oozing with exactly the vitality that you’d expect from your typical government-trained civil servant!  Well, okay, so maybe I’m not exactly Milton Waddams.  All personal resemblance and charm aside, I don’t actually have quite that much paper clutter in my office.  No.  Instead, I just have electronic clutter.  Indeed, if someone spent the time to actually print out all the stuff on my computer here at work, then we wouldn’t need a rocket to get beyond the moon.  We could just build a staircase from the resulting gargantuan pile of paper. 

Why is this the case?  It could be that I’m just a data hoarder.  Some people hoard clothes (that’s not me).  Some people hoard automobiles (that’s not me).  Some people hoard books (okay, that is me).  And some people hoard data.  All data.  All of the way back to class notes from Aerodynamics I (AERSP 311, junior year, professor Bill Holl — great teacher, great guy).  Yes, okay, so maybe I’m a little guilty of all that.  In my defense, however, I will simply say that there’s a lot that goes into making rocket engines and much of it is quite removed from the exciting cutting-metal stuff or the making-smoke-and-fire stuff.  And I get to stick my nose into much of it.  For this article, I am going to reveal super-duper, deep-and-dark secrets that they won’t even teach you in college.  I am going to tell you a little about …

…wait for it…

…project management.  Yes, it’s true:  I am going to invite you into our little piece of office space and I will do so to tell you about how the office has recently evolved.  Also, towards the end as a reward, I’ll give you an update of J-2X testing progress to date.

So, let me introduce you to LEO —

 

No, none of those LEOs.  Instead, let me introduce you to the Space Launch System (SLS) Program Liquid Engines Office (LEO).  The LEO is responsible for development and delivery of the liquid rocket engines for the core stage and upper stage of the SLS vehicle.  For more information about the SLS Program in general, I highly recommend the following site: https://www.nasa.gov/exploration/systems/sls/

(Oh, and I want to say something briefly about the term “liquid engines.”  Probably every single person who would bother to read a blog about engines or rockets or space travel in general knows that this is just a shorthand term.  No, we are not talking about engines made of liquid — although that would be really cool.  “Liquid engines” is a quick and easy way to denote “liquid-propellant rocket engines.”  In case I’ve disappointed anyone, I’m sorry.  If ever we are able to make an engine out of liquid, I promise to be the first to report it.  Probably the most far-out thing that I once heard was the suggestion to make a hybrid rocket motor using solid hydrogen and liquid oxygen.  I cannot even imagine what the infrastructure would be to make the use of solid hydrogen plausible, but you never know…)

For the SLS vehicle, the upper-stage engine is the rocket engine so near and dear to our hearts after several years of design and development and fabrication and assembly and test:  J-2X.  The core-stage engine is the RS-25.  No, the RS-25 is not a brand new engine.  Rather, it is the generic name for that workhorse of the last thirty years, the Space Shuttle Main Engine (SSME). 

At the end of the Space Shuttle Program, there were fourteen SSMEs that had flown in space on the Shuttle and that still had usable life remaining.  I’m not sure that everyone knows this, but rocket engines have limited useful lives.  I guess that most things do, but with rocket engines it’s often pretty short.  Think of them like cherry blossoms (popular motif in Japanese tattoos): amazingly beautiful and quickly gone too soon.  The stresses within an operating rocket engine are tremendous.  For example, the J-2X has an official, useful life of only four starts and less than 2,000 seconds of operational run time after the engine has been delivered for use as part of the vehicle.  No, the engine doesn’t crumble into dust after that, but based upon our certification strategy and on our analysis of margins, that is the official life for our human-rated launch system.  After that point, depending on the proposed usage and risk considerations, and based on the likely reassessment of our margins with the proverbial “sharper pencil,” we can and do routinely talk ourselves into longer active lives for engine hardware.  On the test stand, we can test the J-2X upwards of 30 times and for lots of run time, but that is a lower risk situation.  Nobody is riding the test stand into space. 

Thus, when you come to the end of a program and you have fourteen engines with remaining, usable life, then you’ve got one heck of a residual resource.  In addition, there was one SSME assembled and ready to go, but it never made it to the test stand or the vehicle.  So it’s brand new.  And, on top of that, there were enough leftover pieces and parts lying around of flight-quality hardware to cobble together yet another engine.  And, there’s more! (Yes, I feel like the late-night infomercial guy, “and if you call in the next 10 minutes you will get this special gift!”)  There are also two development SSMEs.  These are not new enough to fly, but they are useful for ground testing and issues resolution.  That means that there are a total of sixteen RS-25 flight engines and two RS-25 development engines available to support the SLS Program. 

However, before your excitement bubbles over, you have to understand that when you see a sign for “free puppies,” you probably shouldn’t take that whole notion of “free” too literally.  As in, well, not at all.  Yes, we still have an extraordinary asset in the residual RS-25 engines.  No question.  But, we have work to do to integrate them into the SLS Program.  In a future article, I will discuss the multiple facets of this work.  By the way, I cannot claim to be immune from the “free puppies” thing myself.  Meet Ruugie –

The Liquid Engines Office (LEO) was formed to manage both the J-2X and the RS-25.  This office will also manage other liquid rocket engines used to support the SLS Program as it matures.  It was decided from a project management perspective that it would be best to have one office manage both engines.  In this way, we can be more efficient by leveraging the expertise across various disciplines and components.  For example, do we really need two turbomachinery subsystem managers?  No, Gary Genge is our turbomachinery subsystem manager and in that position he can understand and evaluate the relative programmatic and technical risk across all of the various turbomachinery pieces under his purview.  If in some utopian future our office responsibilities expands to three or four or eight different engine development or production efforts, we would, in theory, maintain the same structure but provide Gary with the support necessary to effectively manage turbomachinery across so many activities. 

So, for LEO we have subsystem managers for Engine Systems (effectively systems engineering and integration), Engine Assembly and Test (also includes asset management, logistics, and operations), Engine System Integration and Hardware, Valves and Actuators, Engine Control Avionics, Turbomachinery, and Combustion Devices.  LEO is supported by a Chief Engineer, a Chief Safety and Mission Assurance (S&MA) Officer, Program Planning and Control (i.e., the business office), and Procurement.  Plus, of course, we have support from the engineering and S&MA organizations across the many technical disciplines.  The structure is really quite similar to how we’ve been managing J-2X for these past several years.  We’ve just expanded our responsibilities.

So, that’s LEO and I’ll be talking more about RS-25 and SLS in the future.

Now, while I’ve been off doing my little part to get the foundation of LEO solid, including refreshing and getting into place our prime contracts for both J-2X and RS-25, how has J-2X been doing?  Well, in short, J-2X has been just cruising along.  E10002 has gone through six tests on NASA Stennis Space Center (SSC) test stand A-2.  Below are a series of images showing what an E10002 start looks like if you stood in view of the flame bucket (which I would very strongly advise against, by the way):

First, all you see is the facility water being pumped into the flame bucket.  Then you can see the ignition and everything glows orange.  Then the whole flame bucket is filled with exhaust.  And, finally, the exhaust coming barrelling down the spillway and eventually engulfs the camera.  The final step is not shown since there’s nothing to see but solid whitish grayness.

Here are the stats on the six tests:
     • Test:          A2J022          2/15/2013          35 seconds duration
     • Test:          A2J023          2/27/2013          550 seconds duration
     • Test:          A2J024          3/07/2013          560 seconds duration
     • Test:          A2J025          3/19/2013          425 seconds duration
     • Test:          A2J026          4/04/2013          570 seconds duration
     • Test:          A2J027          4/17/2013          16 seconds duration

So the total accumulated time is 2,156 seconds.  Tests #22, #25, and #27 all experienced early cuts, but all three were instigated by different flavors of instrumentation or monitoring system issues or oddities.  The engine is fine and running well.  Some of the key objectives included gathering additional data about the nozzle extension cooling characteristics, additional samples of the turbomachinery design, and main chamber combustion stability trials.  Something else that we did for this test series is that we tested a very special fuel turbopump port cover.  Here’s a picture of it:

Now, port covers are not something about which one usually says anything at all.  What makes this one special is that it was made by using a process known as Selective Laser Melting (SLM).  That is a fabrication method that is somewhat analogous to “3-D printing.”  A long time ago, I wrote a blog article about a gas generator discharge duct that we made for component-level testing using this technique.  This, however, is an engine test and this small, seemingly innocuous, piece of engine hardware may be the humble harbinger of a revolution in rocket engine fabrication.  The fact that we systematically stepped through the process of validating this port cover as a piece of hardware for an engine hot fire demonstration paves the way for pursuing other parts in the future, more complex parts, and, hopefully one day, regular production parts as part of a human-rated launch architecture. 

E10002 was removed from NASA SSC test stand A-2 on April 30th.  It is currently being retrofitted with instrumented inlet ducts and other hardware in preparation for the next phase of testing that will occur on NASA SSC test stand A-1.  As you’ll remember, in the past A-1 was used for the PowerPack Assembly testing.  Well, the talented and productive folks at NASA SSC remodeled the stand back to the configuration for engine testing.  The current plan is to install E10002 into A-1 by mid-May and to perform a series of five to seven tests through probably August.  The reason for using A-1 for the next series is because that stand does not have a diffuser.  That means that we can gimbal the engine, i.e., twist it around as if we were providing steering for a vehicle.  The thrust vector control (TVC) system composed of the hydraulic push-pull actuators that will be performing the gimballing is a component belonging to the stages element of the vehicle.  This testing will be providing those folks with data to inform their system design for the SLS Program.  See, it’s all win-win when we play nicely together.

And, finally, right on the heels of E10002, the assembly of E10003 will commence in June with scheduled installation into NASA SSC A-2 in September.  That’s my report for where things stand.  To finish up, I’ll leave you with a purely gratuitous glamor shot of the J-2X.  Isn’t she pretty?


J-2X Progress: Current Status, The End of 2012

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Once upon a time, not that long ago, people used to communicate by what were known as “letters.”  These were written documents.  Yes, actual hardcopy, paper items. And they were often transcribed by hand or, sometimes, generated on what was known as a “typewriter,” which was basically a manual, analog printer with no I/O port beyond direct keypad entry.  These “letters” were sent to their intended recipients using a small denomination currency with an adhesive backing that is recognized for exchange by only one quasi-governmental agency. 


I know that some of you may have doubts that people communicated with each other in primitive ways prior to email and text messages, but witness the cultural clues from the 1961 song illustrated above. 

It was always believed that the toughest letter to receive was the dreaded “Dear John” letter (as in, “Dear John, I’ve fallen in love with someone else…”).   However, I think t’at the hardest letter to write is the “it’s been awhile” letter.  This one starts, “Well, it’s been awhile since I’ve written.  Sorry.”  This blog article is just like one of those letters.  It’s been awhile since I’ve written one of these articles and I’m sorry about that.  I could give you a big long list of all the really, really serious stuff that I’ve been doing instead, but that’s just a bunch of feeble excuses so I’ll keep them to myself.  Instead, I’ll just get down to business and give you a status report on the J-2X development effort.

Engine #1 (E10001) Testing is Complete!
Over fourteen months and across the span of twenty-one tests, more than 2,700 seconds of engine run time was accumulated and recorded, including nearly 1,700 seconds of hot fire with an instrumented nozzle extension.  With this engine we achieved stable 100% power level operation by the fourth test and full mission duration by the eighth test.  While we don’t have any official statistics on the issue, most folks around here believe that we accomplished those milestones faster than has ever been done on a newly developed engine.  We learned how to calibrate the engine and the sensitivities that the engine has to different calibration settings, i.e., orifice sizes and valve positions.  We were able to estimate performance parameters for the full-configuration of the engine at vacuum conditions and the calculations suggest strongly that all requirements are met by this design and met with substantial margin.  This is significant considering that we’ve long considered our performance goals to be pretty aggressive.  Well, our little-engine-that-could showed us that it did just fine with those goals, thank you very much.

One of the truly unique and successful aspects of the E10001 testing was the testing of a nozzle extension.  This component is a key feature that allows J-2X performance to far exceed that of the J-2 engine from the Apollo Program era.  While it is true that we cannot test the full-length nozzle extension without a test stand that actively simulates altitude conditions, we did test a highly instrumented “stub” version that allowed us to characterize the thermal environments to which the nozzle is exposed during engine hot fire and it demonstrated the effectiveness and durability of the emissivity coating that was used.  This stub-nozzle configuration is actually the current baseline for the in-development Space Launch System vehicle upper stage.

Another key success for E10001 was the demonstration of both primary and secondary power levels with starts and shutdowns from each power level and with smooth in-run transitions back and forth between them.  That smoothness was thanks, in part, to demonstrating our understanding of the control of the engine.  From the very first test it was clear that we understood pretty well how to control the engine in terms of proper control orifices for the various operating conditions.  What we did not entirely understand — in other words the fine-tuning details — we successfully learned via trial-and-error throughout the E10001 test series.  All of this learning has been fed back into further anchoring our analytical tools and models so that we can move forward with J-2X development with a great deal of confidence.

Okay, so that’s a brief description of just some of the good stuff.  We had lots and lots of good stuff with the E10001 testing, far more than just that I’ve discussed here (see previous blog articles).  The somewhat unfortunate part was the way in which the E10001 test series came to an end.  On test A2J021, we had a disconnection between the intent for test and the detailed planning that led to the actual hardware configuration we ran for the test.  That disconnection led to an ill-fated situation.  Let me explain…

The J-2X gas generator has ports into which solid propellant igniters are installed.  These igniters are like really high-powered Estes® rocket motors that light off when supplied with a high-energy electrical pulse.  The flame from the igniter lights the fire of the hydrogen-oxygen mixture during the engine start sequence.  It’s essentially the kindling for the fire of mainstage operation.  The igniters perform this function at a very specific time during this sequence.  If you try to light the fire too early, then you may not have enough propellant available in a combustible mixture so you get a sputtering fire.  If you try to light too late, then you may have too much propellant built up such that rather than getting a good fire, you get an explosion instead.  But here’s a key fact: You have to plug them in or they don’t work.

Have you ever stuck bread in the toaster, pushed down the plunger, gone off to make the coffee, and come back only to find that your darn toaster is broken?  You curse a little because you’re already late for work and this is the last darn thing you need.  You would think that somebody somewhere could make a toaster that lasts more than six months or a year or whatever.  For goodness sake!  We put a man on the moon and yet we can’t … oh, wait … um … ooops, it’s not plugged in.  My bad.

In a nutshell, that’s what happened on test A2J021.  The electronic ignition system sent the necessary pulse, but because of the uniqueness of our testing configuration as opposed to our flight configuration the wires carrying the pulse weren’t hooked up to the little solid propellant igniters in the gas generator.  In the picture below you can see the external indication that something was not entirely good immediately after the test.  The internal damage was more extensive to both the gas generator and the fuel turbopump turbine.

Many years ago, I met an elderly engineer who was still on the job well into his 80’s because he loved his work.  His entire career had been dedicated to testing.  He’d actually been there, out in the desert, in the 1940’s testing our very earliest rockets as part of the Hermes Project.  One day, they had a mini disaster on the launch pad.  He told me that the rocket basically just blew up where it sat.  Boom and then a mess.  And, it was his job to assemble the test report.  Being a conscientious, ambitious, young engineer, he recorded the facts and offered a narrative abstract and extensive, annotated introduction that categorized the test as, well, a failure.  Not long after submitting his report, one of the senior German engineers in the camp came into his office, put the test report down on the desk, and said that the tone of the report was entirely wrong.  He said, “Every test report should begin with: ‘This test was a success because…'”  The purpose of testing is to gather data and learn.  If you learn something, then your test was, by definition, a success on some level.  I’ve tried very hard to remember this very important bit of wisdom.

So, A2J021 was a success because we learned that we had some deficiencies in our pre-test checkout procedures.  It was a success because it was an extraordinary stress test on the gas generator system.  No, it didn’t recover and function properly, but neither did the engine come apart.  While that might seem like a minor detail, when you’re hundred miles from the surface of the earth, you would much rather have a situation where an abort is possible than a failure that could result in collateral vehicle damage and make safe abort impossible.  We have a stout design.  Good.  Also, this test failure was due to a unique ground test configuration.  In flight, it’s not really plausible just because we would never fly in this configuration.

So, E10001 completed its test program with a bang.  Kinda, sorta literally.  But it was nearly the end of its design life anyway, so we didn’t lose too many test opportunities, and, as I said, even with test A2J021 the way it happened we learned a great deal.  Overall, the E10001 test series was an outrageous success.  Rocketdyne, the J-2X contractor, ought to be darn proud and so should the outstanding assembly and test crews at the NASA Stennis Space Center and our data analysts here at the NASA Marshall Space Flight Center.  Bravo guys!  Go J-2X!

Power-Pack Assembly 2 (PPA-2) Testing is Complete!
Over ten months and across the span of thirteen tests, nearly 6,200 seconds of engine run time was accumulated and recorded on the J-2X Power Pack Assembly 2.  That’s over 100 minutes of hot fire.  Three of the tests were over 20 minutes long (plus one that clocked in at 19 minutes) and these represent the longest tests ever conducted at the NASA Stennis Space Center A-complex.  But more than just length, it was the extraordinary complexity of the test profiles that truly sets the PPA-2 testing apart.

Because PPA-2 was not a full engine with the constraints imposed by the need to feed a stable main combustion chamber, and because we used electro-mechanical actuators on the engine-side valves and hydraulic actuators on the facility side valves, we could push the PPA-2 turbomachinery across broad ranges of operating conditions.  These ranges represented extremes in boundary conditions and extremes in engine conditions and performance.  On several occasions we intentionally searched out conditions that would result in a test cut just so that we could better understand our margins.  As the saying goes: It’s only when you go too far do you truly learn just how far you can go.  We successfully (and safely) applied that adage several times.  In short, we gathered enough information to keep the turbomachinery and rotordynamics folks blissfully buried in data for months and months to come. 

On an interesting and instructive side note, the PPA-2 testing also showed us that we needed to redesign a seal internal to the hydrogen turbopump.  In the oxygen turbopump, you have an actively purged seal between the turbine side and the pump side.  After all, during operation you have hydrogen-rich hot gas pushing through the turbine side and liquid oxygen going through the pump side.  You obviously don’t want them to mix or the result could be catastrophic.  That’s why we have a purged seal.  But for the hydrogen turbopump you don’t have such an issue.  During operation, at worst should the two sides mix you could get some leakage of hydrogen from the pump side into the turbine side that is already hydrogen rich.  In order to maintain machine efficiency, you don’t want too much leakage, but a little is not catastrophic (and can be used constructively to cool the bearings).  What could be dangerous at the vehicle level, however, is if you have too much hydrogen floating around prior to liftoff.  This is especially true for an upper-stage engine like J-2X that’s typically sitting within an enclosed space until stage separation during the mission.  You could have the engine sitting on the pad for hours chilling down and filling the cryogenic systems and you don’t want gobs and gobs of hydrogen leaking through the turbopump since any leakage ends up within the closed vehicle compartment housing the engine.  That’s just asking for an explosion and a bad day.

To avoid this, within the J-2X hydrogen turbopump we have what is called a lift-off seal.  And, as the name applies, it’s a seal that actively lifts off when we’re ready to run the engine.  When the engine is just sitting there chilling down, not running, with liquid hydrogen filling the pump end of the hydrogen turbopump, the seal is, well, sealed.  Then, when we’re ready to go, it unseals and allows the turbopump to operate nominally.

During the PPA-2 test series we found that we formed a small material failure within the actuation pieces for our lift-off seal.  Then, upon analysis of the test data and a reassessment of the design, we figured out what was most likely the cause and Rocketdyne proposed a redesign to mitigate the issue.  Again, going back to that important piece of wisdom: This testing was a success because, in part, we learned that we needed a slight redesign of the lift-off seal.  That’s the whole purpose of development testing!  Everything always looks great when it’s just in blueprints.  It’s not until you hit the test stand do you truly learn what’s good and what need to be reconsidered.  In the end, this sort of rigor and perseverance is what gives you a final product that you feel good about putting in a vehicle carrying humans in space.  And that, truly, is what it’s all about.

As with E10001, the PPA-2 test series was simply an outrageous success.  Rocketdyne should be proud and so should the outstanding assembly and test crews at the NASA Stennis Space Center and the data analysts at the NASA Marshall Space Flight Center.  Bravo guys!  Go J-2X!

Engine #2 (E10002) Assembly is Underway
Our next star on the horizon is J-2X development Engine 10002.  It is being assembled right now, as I’m typing this article.  It is slated for assembly completion in January 2013 and it will be making lots of noise and very hot steam in the test stand soon after that.  While our current plans are to first test E10002 in test stand A2, we will later be moving it to test stand A1.  This, then, will be the first engine then to see both test stands.  The more important reason for the A1 testing, however, is because that will give us the opportunity to hook up some big hydraulic actuators and gimbal the engine, i.e., make it rock and tilt as though it were being used to steer a vehicle.  Now that will be some exciting video to post to the blog!  I can’t wait.

 
Happy New Year!
So, this has been my “it’s been awhile” letter.  Hopefully this will bring everyone up to speed with where we stand with J-2X development.  In my next article, I will share with you some of what’s been keeping me from my J-2X article writing over the last several months.  And, hopefully, it won’t be several months in the making.  So, farewell for now and Happy New Year!  On to 2013 and another great year full of J-2X successes.  Go J-2X!

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

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In the last article, we talked about the measurement of propellant flow during a test.  Propellant is the stuff we put into the rocket engine.  What we get out of the rocket engine is thrust.  We get propulsion.  Or, in the immortal words of Salt-n-Pepa, 1987, we get push… “Push it real good.”

But how do you measure “push,” or in other words, force?  The simple answer can be found through one of the most frightening household appliances any of us own:  The dreaded bathroom scale…

A bathroom scale works by pushing back at your weight when you stand on it.  Your weight is a force caused by your body mass and the Earth’s characteristic acceleration of gravity.  The scale pushes back with a spring system but deflects slightly under the load.  The scale is then measuring the deflection allowed by the springs at the equilibrium where the spring force exactly counteracts your weight.  More weight pushing down results in more deflection to the equilibrium point and that thereby results in a bigger reading (i.e., think: the Monday after Thanksgiving).

The way that we measure rocket engine thrust is basically the same thing except that instead of measuring something between zero and — as in the bathroom scale picture above — 300 pounds, we’re measuring hundreds of thousands of pounds of force.  Or, in the case of very large engines like the F-1 or the RD-170, we’re measuring over a million pounds of thrust.  That requires a system just a bit more rugged even if the principles remain the same.  What we use rather than bathroom scale and springs are things called “load cells.”  Below is an example of a generic load cell design:

The gray object with the funky cut through it is a metal piece.  As you can imagine, when forces are applied as shown, that slot on the right-hand side will tend to close slightly.  In turn, that would cause the metal on the left-hand side to stretch slightly as the whole thing bends a very little amount.  We measure that slight stretching of the metal on the left-hand side with a strain gauge bonded to the surface of the metal.  A strain gauge is a small electrical device that changes resistance when stretched.  Using an electrical circuit known as a Wheatstone Bridge, we measure small changes in electrical resistance caused by the slight deformations of the load cell.  The amount of stretch can be astonishingly small and yet good strain gauges and good electrical interpretation of the output can yield very accurate data.  The load cell is then calibrated using a known applied load and measuring the resulting strain (i.e., metal stretch).  You now have a more rugged version of a bathroom scale.  Apply a load, get a reading, and, ta-da, you’ve measured push.

Actual load cells used for rocket engines can take different forms from the generic cell shown here.  Any way that you can get an applied load to result in a slight, measureable stretching of metal (while obviously avoiding yielding or buckling) is a valid load cell design.

Above is a picture of a vertical load cell arrangement on test stand A-2 at the NASA Stennis Space Center, where we’ve been testing J-2X development engine E10001.  There are two pieces in series.  The bottom piece, the big chunk of metal with a bunch of crazy holes and slashing cuts through it, is called a “flexure,” which, to me, seems to be a silly name since it doesn’t look very flexible at all.  What it does, however, is effectively make sure that the load entering the load cell is properly directed through the intended vector.  Any skewing of the input off from the intended axis and your results could be erroneous.  The brown cylindrical thing above the flexure is the actual load cell.  You can see the strain gauge wires coming out of it that are fed into the data acquisition system.  This two-piece combination is effectively analogous to a spring in your bathroom scale.

The next item to discuss is how you put load cells into the structure of the test stand so that they can do their job.  On a bathroom scale, the thing that you step onto is essentially a platform “floating” above the base.  It has to be free to move so that the springs can compress honestly.  If there was some interference with this movement, then the reading would be wrong.  The same is true on the test stand when measuring engine thrust.  It is necessary to use a free-floating platform.  The picture below is a drawing of the platform used on test stand A-2.

The engine has a single input point as shown — the gimbal bearing that we’ve discussed before in previous articles — and there are three load cells above the platform.  This is not the only possible way to do it.  Other test stands use a rectangular pattern of four cells.  Or, if it’s a smaller system, you might be able to use just a single load cell.  The important point is that the load cells are in between the pushing engine and the resisting test stand.  Put into the structure of the test stand, and viewed from the side as in the picture below, you can see the whole stack up.  On the bottom is where you attach the engine.  In the middle is the platform into which the engine pushes.  And then the platform is connected to the structure of the stand through the load cells.  The structure of the stand has to be strong enough to absorb the thrust of the engine without distorting.  It has to be fixed (the mythical “immovable object” from physics class).  So, as you can imagine, when you’re talking about hundreds of thousands or even millions of pounds of force, the test stand structure is pretty darn stout.  We usually refer to the primary structure responsible for resisting the force of the engine as the “thrust take-out” structure. 

A final subject to mention is that matter of tares.  Tares are corrections to measured data.  For example, when you step on the bathroom scale and you’re dressed, do you subtract off an estimate for the weight of your clothing and your shoes?  If so, then you’re making a correction for a particular tare.  Of course, you have to do this accurately (honestly).  If I assume that my clothes and shoes are made of lead, for instance, then I can declare that I weigh the same as when Salt-n-Pepa were releasing their first albums.  But that’s not quite the truth.  Getting your tares correct is important for interpreting your data correctly.

When measuring rocket engine thrust you have lots and lots of corrections to the raw data that you measure with your load cells.  This is because, in truth, the gimbal bearing is not the only connection between the engine and the test stand.  While you’d really like to have that perfectly free-floating platform situation, you’ve got to have, for example, propellant feedlines hooked up to the engine.  Flexible bellows are built into the line so that they’re not completely stiff and thereby interfering with the movement of the platform, but they still absorb some of the thrust load and, therefore, make the raw thrust reading skewed.  There are a number of other such corrections that need to be made such that the whole calculation process related to tares can get a bit cumbersome with all its many pieces, but nobody ever said that developing rocket engines was supposed to be easy, right?

Now, between this article and the previous one, you have a good idea of how we get basic performance data from rocket engine testing and also the necessary configuration of the test stands that allow us to gather this information.  The smoke and fire and rumbling roar of an engine test is all very impressive, but for us Datadogs, it’s the data that matters most.  We get lots and lots of data from every test, but propellant flow rates and engine thrust are the most important in terms of understanding how an engine fits into a vehicle and a mission.

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

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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!

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.

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. 

 

J-2X Progress: Once Upon a Time at Stennis…

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I enjoy movies.  I don’t get to watch much television due to other endeavors that consume much of my time, but if I do it’ll almost always be one of four things on the screen:  some news program, a sporting event, a history program, or a movie.  And I like lots of different kinds of movies.  Some of my favorites include: The Hustler, Singing in the Rain, Rocky, Schindler’s List, Barfly, Hannah and Her Sisters, Fargo, The Apartment, The Godfather, Leaving Las Vegas, The Deer Hunter, Hoosiers, Nobody’s Fool (the Paul Newman one).  I don’t believe that one could decipher a pattern from that list other than the fact they all follow the classic narrative structure:

Think of the classic “stranger comes to town” story.  (1) It’s a quiet little town and all is peaceful.  (2) Then a stranger comes to town and stirs up all kinds of trouble.  (3a) In the end, the stranger marries and settles down with the prom queen and everyone learns to live with one another.  Or, (3b) in the end, the stranger ends up mysteriously dead and lying in the gutter along the road leading out of town and they secretly bury him promising never to mention it to anyone from out of town.  Or, (3c) in the end, the stranger ends up mayor of the town by exposing and driving out the secretly corrupt sheriff.  Obviously, the possibilities are endless and that’s why there are thousands and thousands of stories to be told.  But the root of all of this is the middle block, “…something disturbs that situation and troubles ensue…”  Nobody ever tells an interesting story where nothing happens.  And with no “troubles” of some sort, nobody cares about the resolution.

So, that brings me to rocket engine testing and the fact that it is always interesting.  This article is intended to bring you up to date on the status of our J-2X test campaign at the NASA Stennis Space Center in southern Mississippi.  Remember, we last left our heroes on test stand A1 with PowerPack-2…

Test A1J015, J-2X PowerPack-2: It ran 340 seconds of a planned 655 seconds duration.  The test profile called for simulated primary mode and secondary mode (i.e., throttled) operation.  Also, throughout the test, turbomachinery speed sweeps were planned meaning that we systematically varied turbine power, increasing and decreasing, to force the pumps through a broad range of conditions.  It was during one of these sweeps that the fuel turbopump crossed a minimum speed redline and the test was cut short.  Before the test, we knew that it would be close and the analytical prediction was just enough off from reality to cause the early cut.  Nevertheless, most of the primary objectives were achieved and the test was a success. 

One of the things that we often talk about when discussing an engine test is the “test profile” or sometimes the “thrust profile.”  The test/thrust profile is the plan for what you’re going to do during the test.  When we say that we had a planned duration of 655 seconds, that value comes from the test profile that is agreed upon prior to the test.  Usually a test/thrust profile is a single page showing engine power levels and propellant inlet conditions, but for these complex PPA-2 tests, the test profile can be expanded to include such things as these turbomachinery speed sweeps.  To give you an idea of what an engine test/thrust profile looks like, here is one for a Space Shuttle Main Engine (SSME) test performed back in 2001.  It contains a wealth of knowledge about the test to be run.

Test A1J016, J-2X PowerPack-2:  It ran 32 seconds of a planned 1,130 seconds duration.  In this case, unlike the previous test, because we cut so early we can’t really say that it was mostly a success.  However, every time that you chill an engine, successfully get it started, and shut it down safely, you have accomplished something significant and you are always collecting data and learning.  The early cut in this case had nothing to do with the PowerPack-2 performance.  Rather, it was a facility issue, a hydrogen fire due to a leak.  As I’ve said before, the PowerPack-2 is an oddball test article in that it is half engine and half facility.  That makes the interfaces technically difficult in some cases due to thermal and structural loads.  The leak and fire in this case was on the facility side near one of these difficult interfaces. 

Below is a picture captured off a video taken during the test and behind the structure and the piping you can see the bright orange flame that resulted in the early cut.  This issue of hydrogen leaks and fires has been somewhat recurring so a team of NASA and contractor folks stepped forward to work towards a resolution of the issue.

Test A1J017, J-2X PowerPack-2:  It ran the full, planned 1,150 seconds duration.  That’s over 19 minutes of continuous rocket engine operation and that’s pretty amazing.  It was the longest, most complex engine test ever conducted across the long history of the NASA Stennis Space Center A Complex.  We did some wacky stuff on test stand A1 during the XRS-2200 (linear aerospike engine) development effort and there were a couple of longer SSME tests in the B Complex twenty-some years ago, but test A1J017 stands out for the combination of complexity and duration.  The test profile contained over a dozen unique, steady state “set points,” i.e., prearranged combinations of engine operational conditions and facility boundary conditions.  The objectives of this test included speed sweeps for the oxidizer turbopump and an examination of cavitation performance for both the oxidizer pump and the fuel pump.  Pulling off this test was a dazzling success with many people deserving credit.

So, trouble ensues (hydrogen fire on test #16) and the combined team of NASA Stennis, NASA Marshall, test operations and support contractors, and Rocketdyne worked through to a resolution of the issue and a new situation of unprecedented success has been achieved.  It’s easy to write a blog like this when reality lines up so conveniently in the narrative form. 

But back at the ranch, our heroes find J-2X development engine E10001 on test stand A2…

To refresh your memory, we’d last tested E10001 on stand A2 back in December of last year.  Back then, we were testing the engine without a nozzle extension and not using the passive diffuser system on the stand.  This year, we were going to get back to testing E10001 but now with a nozzle extension so that necessitated use of the passive diffuser.  The Stennis folks installed a clamshell and seal apparatus that connects the engine to the diffuser thereby allowing the diffuser to “suck down” to pressures lower than sea level ambient.  In my crude sketch below, I try to show you how this fits together.

A key piece in this arrangement is the clamshell seal.  Whereas the engine is obviously metal and the clamshell and diffuser and big pieces of structural metal, the clamshell seal is a fibrous/rubber-ish piece that has to provide the seal that allows the whole thing to work together and simulate altitude operation when the engine is running.  It has to be strong yet compliant so as to accommodate movements of the nozzle during hot fire.  To give you an idea of how strong it needs to be, let’s calculate the force imposed on the seal during operation.  Ambient sea level pressure is 14.7 psia (pounds per square inch, absolute).  Let’s say that in the diffuser, during operation, it will be about 10 psi lower than sea level ambient.  In reality, the pressure will be slightly lower than that, but 10 is a nice round number to work with.  Let’s further say that the diameter of the nozzle at which the seal is attached is about five feet (or, 60 inches).  That’s pretty close to reality, give or take a bit.  And, let’s say that the seal itself is about six inches in width.  So, the total area of the seal is:

So, if the pressure differential across the seal is 10 pounds per square inch and you have 1,244 square inches of surface area, then that makes for over 12,000 pounds of force — or more than 6 tons!  Wow, so that seal and the brackets that holds it in place still needs to be pretty darn tough.

Test A2J011, J-2X E10001: It ran 3 seconds of a planned 7 seconds duration.  The early cut was due to a facility redline violation; specifically, the measured pressure within the clamshell did not drop down the way that it was supposed to.  Post-test inspections quickly revealed why this redline violation occurred.  The clamshell seal was torn up.  If the seal doesn’t seal, then the pressure differential is not maintained and so, appropriately, we tripped a redline.

An informal team was assembled of NASA, contractor, and Rocketdyne folks and the design deficiency was quickly identified.  New parts were designed and fabricated and, in a matter of just a couple of weeks, we were once again ready for test.

Test A2J012, J-2X E10001:  It ran the full, planned 7 seconds duration.  The objectives for this test were to demonstrate that the clamshell, seal, and diffuser arrangement was properly working and to perform a bomb test in the main chamber.  The testing arrangement worked perfectly and the bomb test did not reveal any combustion stability issues. 

Test A2J013, J-2X E10001:  It ran the full, planned 40 seconds duration.  This was yet another bomb test and again there was no combustion stability issue uncovered.  The neato thing on this test was that while the engine started to primary mode operation (i.e., 100% throttle), it switched to secondary mode operation (i.e., throttled) mid-test.  This was the first operation of the complete J-2X engine (as opposed to just the powerpack portions) in secondary mode. 

Test A2J014, J-2X E10001:  It ran the full, planned 260 seconds duration.  This test represented several more “firsts” for J-2X.  This was the first time that the J-2X was started directly to secondary mode.  It was the first time that the J-2X switched, in run, from secondary mode to primary mode.  This was the first J-2X test with a stub nozzle extension that offered the opportunity to perform an in-run calibration of the facility flow meters and, in so doing, provide for a good estimation of engine performance.  It turns out that E10001 is, to our best understanding, exceeding expectations in terms of required performance.

Again, the old narrative structure holds:  New guy comes to town (the stub nozzle extension).  The situation changes (new test stand configuration to accommodate the stub).  Troubles ensue (the clamshell seal gets torn up).  Resolution is found (new design for clamshell seal attachments).  And a new situation is achieved (we’re knocking off successful test after successful test). 

But, there is a twist (literally) to our denouement.  I’ll explain this twist by starting with a picture:

Can you see it?  This is a picture of the fuel inlet duct.  Remember, this duct has an inner and an outer shell (or bellows as we call them) so that in between there will be vacuum to keep the hydrogen cold, like a Thermos® bottle.  Between tests, one of the customary inspection techniques used to ensure that you’re good to go for the next test is to do a series of helium leak checks.  You systematically pressurize different portions of the engine and make sure that everything is still sealed up tight.  Well, when they pressurized this portion of E10001, they got what we’re calling “squirm.”  If you look closely at the duct you’ll see that on the left-hand side the convolutions are bunched together and on the right-hand side they’re spread apart.  This indicated that there was leak in the inner bellows of the duct so that the cavity between the two bellows was pressurizing with the leak-check helium.  The squirm effect was due to the outer shell was deforming — squirming — due to that pressurization of the vacuum cavity. 

Now, there are several important things to note about this.  First, this particular duct is a heritage piece of hardware.  It was not made for E10001.  It was made during the Apollo era for J-2 and J-2S, forty years ago.  It had seen its fair share of hot-fire history long before it reached E10001.  Second, the new ducts being built for J-2X have a design modification that ought to mitigate this kind of failure.  Third, we can see in the test data, with perfect hindsight, exactly when the leak occurred in test A2J014 and the engine ran for some time with the leak and nothing catastrophic happened.  Thus, while nobody is happy when something breaks, in this case there’s no need for overreaction.

Getting back to the narrative structure and this little twist at the end, I kind of think of this like a teaser — a cliff-hanger — that leads to a sequel.  Will our intrepid heroes dig their way out of this situation?  Will the test program recover and move ahead to new successes and glory?  Or will the monster creep up from the dark, dank Pearl River swamps and terrorize the test crew…?

…oops, wrong movie. 

[Hint:  We’ll be fine.  Already moving out at full speed.  In the immortal words of Journey (i.e., Jonathan Cain, Steve Perry, and Neal Schon) “Oh the movie never ends. It goes on and on and on and on…”]


 

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