Liquid Engines Extra -- Introducing LEO
Posted on May 08, 2013 02:57:33 PM | William D. Greene
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: http://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
Posted on Jan 08, 2013 10:01:38 AM | William D. Greene
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
Posted on Oct 16, 2012 08:30:59 AM | William D. Greene
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
Posted on Sep 13, 2012 11:59:17 AM | William D. Greene
I've had jobs of some flavor almost continuously since I was fourteen years old. From delivering newspapers to cutting grass to flipping burgers to showing movies to mixing giant vats of coleslaw to instructing an aerodynamics laboratory course to hand counting vehicular traffic to researching the derivation of combustion stability equations, before I finally settled into something resembling a career (and while I was winding my way through the maze of secondary and higher education), I did all kinds stuff. But when I did get my first job after graduate school it was doing test data reduction and performance calculations for the Space Shuttle Main Engine (SSME).
How cool is that? Trust me: very cool. I consider myself to be extraordinarily lucky in this regard. This article is going discuss engine performance measurements and so it's going to reach back to my very roots in this business. And that sounds like fun!
In past articles we've discussed rocket engine performance. We've talked about the "Big Three" operational points and performance measures that characterize an engine: thrust, mixture ratio, and specific impulse. Okay, but how do we know what these values are for any given engine? I mean, we can do calculations with analytical equations, formulas, algorithms, or models that tell us what these parameters ought to be for a given rocket engine design; but when I've got an actual rocket engine sitting in front of me -- a big, shiny, complex hunk of metal standing ten feet high and weighing thousands of pounds -- how do I know how it actually functions and performs?
Well, duh, you test it. Of course! But making smoke and fire (and steam) does not, by itself, give you any data. The most that you could say from just watching an engine test is that it's really, really noisy and that it makes a really, really big exhaust plume. So, more than just observe, you have to take measurements during the test. That's how you get data. As I've said many times, there are only two reasons to conduct rocket engine tests: (1) to impress your friends, and (2) to collect data. In order to get data on the "Big Three," you need to measure thrust and you need to measure propellant flowrates. For this article, we're going to focus on propellant flowrates. I will talk about thrust measurement in the next article.
Propellant flows are measured on a rocket engine test stand with "propellant flowmeters." Makes sense, right? But calling something a "meter for flow" doesn’t tell you how it works. That's like saying, "How do I make popcorn?" Answer: "With a popcorn maker." No kidding. Thank you for playing and you've conveyed no useful or interesting information.
There are a number of different ways to measure fluid volumetric flow. The units that we use for the very large flowrates feeding an engine are turbine flowmeters. Have you ever blown into a small fan that's turned off or perhaps a pinwheel? If you blow hard enough, you can make the fan or pinwheel spin. That is, quite simply, how a turbine flowmeter works: it's a fan, i.e., turbine, stuck in a tube that spins as fluid flows through it. The faster the fluid flows, the faster the turbine spins. The thing that we measure is the speed of the turbine spinning. The turbine has a number of blades (just like that small fan that you blow into). We pick a spot on the tube in which the turbine sits and count how many blades pass by. If we count, say, ten blade passes in a second, then there is more flow than if we'd only counted eight blade passes in a second.
So, how do we count blade passes? Well, there's a little window in the side of the propellant duct and we sit a young college co-op in front of the window with a little hand clicker and scream "GO!" from the blockhouse… Okay, I'm fibbing. We don't treat our co-ops nearly that bad. Usually. Besides, there would be no way that the human eye and brain could keep up since we're talking about hundreds of blade passes per second. Instead, we measure it electronically. Each blade contains a magnet in the tip. The sensor on the outside of the tube is activated by the magnet. Each magnet pass generates an electronic pulse or blip -- what we call a "pip" -- and we keep a continuous count of these accumulated pips. The pip count is then recorded with each time step in the data collection process. Then, after the test, we can translate this ever-increasing pip count into a pip-rate based upon these recorded times. Mathematically speaking, the pip-rate at any given point is the slope of the pip-count plotted against time.
In order to translate a pip-rate into a volumetric flowrate, such as gallons per minute (gpm), the flowmeter needs to be calibrated. We need to know how much flow is required to generate a blade passage, i.e., a pip. If, for example, we knew that the passage of one gallon was enough to move the turbine exactly one blade pass of rotation, then a measured 100 pips-per-second would equal 100 gallons-per-second, or 6,000 gpm. Thus, calibration of a flowmeter consists of flowing a known volume of fluid through the meter and counting the pips read:
The truth is that it's a bit more complicated in that the calibration varies with the speed of the turbine due to kinetic and mechanical issues of the rotating hardware and due to fluid dynamic effects of the fluid interacting with the turbine blades. However, these are secondary effects as compared to the simple notion of figuring how much a pip is worth in terms of volume.
Luckily enough (or, really, strategically enough), the engine test stands are themselves set up to function as a calibration facility for the flowmeters. This is because the propellant tanks have a known geometry and are equipped with fluid level sensors.
As shown in the figure above, if we know at a particular time the height of the fluid in the tank and then, at a later time, we know a lower height of the fluid, then, using tank geometry, we know the volume of fluid that exited the tank and ran through the flowmeter. In practice, we actually perform this calibration during an engine test. That way we can be assured that the flowmeter rotor is spinning at a speed representative of where we’ll need measurements.
An observant reader would note here that if we know the volume consumed over time just from the level sensors in the tank, then we don't need a flowmeter in the middle. All you need is volume divided by time, right? The problem is one of fidelity. Because the level sensors are discrete points on the pole submerged in the tank, the measures of volume used for calibration are relatively big chunks, as in enough propellant to run the engine tens of seconds. In order to get a decent calibration across several discrete level sensors, we typically need to run between 100 and 150 seconds of steady, mainstage engine conditions. The use of a calibrated flowmeter allows you to see variations in flowrate at much smaller time increments and this allows us to collect and observe more data with regarding to engine characterization at different conditions. You can almost think of the flowmeter as a useful interpolation tool between large chunks of time and consumed propellant.
You will note that so far we've just talked about volumetric flowrates and yet, when we talk about engine performance we refer to mass flowrates. The difference between the volume and the mass of something is its density. For our very pure propellants, fluid density is simply a function of fluid temperature and static pressure. So, we take temperature and pressure measurements immediately downstream of the flowmeter and, using either an interpolated look-up table or empirical curves, we can get density. So, you put it all together and you end up with something along the lines of the following:
That is how you measure and calculate the mass flowrate of the propellants flowing through the feedlines and going into the engine using a turbine flowmeter. The item from the "Big Three" to which this can be applied directly is the engine inlet mixture ratio, which is defined as the oxidizer mass flowrate divided by the fuel mass flowrate.
However, depending on the engine and vehicle design, not all of the propellants that go into an engine go overboard. Often, warmed propellants are returned from the engine to the stage to act as pressurizing gases for the stage propellant tanks. On the Space Shuttle, both gaseous oxygen and gaseous hydrogen were flowed back to the stage for this purpose. The rocket equation that essentially defines the parameter we know as specific impulse is only concerned with propellants that leave the vehicle so for specific impulse calculations you need to use inlet mass flow minus pressurization flow.
As compared to the engine inlet mass flowrates, which for large rocket engines can amount to hundreds of pounds-mass per second, the pressurization flowrates are typically less than one or two pound per second. Flows this small are more effectively measured using flowmeters different from the turbine flowmeters I've described above. For our engine testing we use Venturi meters for these small flows. Venturi meters use a variable flow area coupled with pressure measurements to feed Bernoulli Equation relationships between pressure and fluid velocity. Once you know the fluid velocity, fluid density, and fluid flow area at any point, you can then calculate mass flowrate (for now, at least, I'll not go any further with Venturi meter calculations).
This, then, wraps up the story with regards to propellant mass flow measurements and calculations on the engine test stands. In the next article, we’ll go into the measurement of and calculation of thrust. All of this discussion reminds me so much of my first days/weeks/months on the job working with SSME test data. At first, it was just a bunch of bewildering numbers and data reduction tools and rules and calibration factors and work procedures. I had no idea what was going on. But gradually, as I dug into the data and talked to people and dissected the computer codes and tools we used, I began to piece it all together as to what these measurements and calculations actually meant. Seemingly every day brought a new epiphany in understanding. Boy oh boy, that was fun!
Inside The J-2X Doghouse: Beyond the Gas Generator Cycle
Posted on Aug 08, 2012 09:19:48 AM | William D. Greene
Okay, I admit it: I'm a sucker for the Olympics. I watch with rapt attention to sporting events that I would otherwise never consider viewing other than under the once-every-four-years heading of the Olympics. Why is that? Perhaps that is somehow a measure of my shallowness as a sports fan. Nevertheless, I was truly on the edge of my seat watching the women's team archery semi-finals and finals. Great drama. Wonderful competitors. Exceptional skills. Bravo ladies!
Another thing that I find fascinating about the Olympics is the fact that it brings together such a broad range of people. No, I’m not going to trail off in a chorus of Kumbayah. You simply cannot deny, however, that during the opening ceremonies you see people of every possible color and shade, from every corner of the planet, straight hair, curly hair, black hair, blonde hair, red hair, eye colors to fill a rainbow, and most startlingly, such an amazing collection of body types. These are all world-class athletes and yet they're often so different from each other. I like seeing the six-foot-seven volleyball player walking next to the four-foot-ten gymnast. I like seeing the contrast of the marathoner and the shot-putter. We're all the same species, but, my goodness, we come in an amazing array of shapes and sizes and various accoutrements.
The Pivot to Topic
Rocket engines, too, come in an array of shapes and sizes and various accoutrements (…I bet that you were wondering when or how I'd turn the conversation on topic). I know that this is a blog dedicated nominally to J-2X development, but I think that it's important to understand where the J-2X fits in this family of rocket engines. So, let's start with a table of top-level engine parameters:
Note that is list is nowhere close to being comprehensive. There are lots and lots of rocket engines out there including those currently in development or in production and many that have been retired (like the F-1A in the table). And if you open the window a little wider to include engines originating from beyond our shores, then you've got many more Soviet/Russian, European, Japanese, and Chinese engines to consider. All I want to do here is expose you to some basic yet significant differences between this small set of examples. Interestingly, if you can understand these few engines, then you can understand most of rest of the ones out there as variations on these basic themes.
Please allow me to introduce you to the engines listed in the table.
• Of course, the J-2X needs no further explanation for anyone who reads this blog regularly.
• The RL10 is a small engine that has been the product of Pratt & Whitney since the late 1950's. Over the past sixty years it's evolved and matured. It was actually used on a NASA vehicle back in the 1960's, the Saturn I launch vehicle upper stage (S-IV). Today it's used, in different variants, as an upper stage and in-space engine for both the Atlas V and Delta IV launch vehicles.
• The RS-25 is another name for the Space Shuttle Main Engine (SSME). The development of the SSME began with research efforts in the late 1960's, using a great deal of knowledge gathered from the development of the original J-2, and it was first tested in 1975 and first flew on STS-1 in 1981. The RS-25 engine is now designated to be the core stage engine for the next generation of launch vehicles under the Space Launch System (SLS) Program.
• The F-1A was an upgraded version of the F-1 engine that powered the first stage (S-IC) of the mighty Saturn V launch vehicle that first took man to the Moon. The F-1A was a more powerful version of the F-1 with a handful of design changes intended to make it cheaper yet more operable and safe.
The Key is in the Power
In a blog article here over a year and a half ago, I introduced you to the gas generator cycle engine. The key philosophical point discussed in that article about what makes a rocket engine an engine is the fact that it feeds and runs itself. It does this by finding a means for providing power to the pumps that move the propellants. The origin for this power is the key to any rocket engine cycle. In a gas generator engine, this power is generated by having a separate little burner that makes high-temperature gases to run turbines that makes the pumps work. Below is a schematic for such a system. You've seen this schematic before and it is very much like J-2X.
MCC = Main Combustion Chamber
GG = Gas Generator
MFV = Main Fuel Valve
MOV = Main Oxidizer Valve
GGFV = Gas Generator Fuel Valve
GGOV = Gas Generator Oxidizer Valve
OTBV = Oxidizer Turbine Bypass Valve
Behold Now Behemoth
The F-1A power cycle is similar to the gas generator cycle shown above in that it is still a gas generator cycle, but rather than two separate turbopump units, there was only a single (huge) unit that contained both pumps. So, a single turbine was used to power both pumps rather than having two separate turbines like J-2X. Going back to the table, you will see that the F-1A was different from the J-2X also in the fact that the propellants were different. The J-2X uses hydrogen for fuel and the F-1A used RP-1 (FYI, RP-1 stands for "rocket propellant #1" and is actually just highly purified, high quality kerosene). The chief difference between hydrogen and kerosene is chemistry. A hydrogen-fuel engine will get higher specific impulse than a kerosene-fuel engine but kerosene engines have the distinct advantage of being able to generate more thrust for a given engine size. With a kerosene engine, you are simply throwing overboard more massive, high-velocity propellants in the form of combustion products. Hydrogen is light and efficient from a "gas mileage" perspective but kerosene gets you lots and lots of oomph. That's why you typically use it for a first stage application like on the Saturn V vehicle. You want to have lots of oomph to get off the ground. Later, on the upper stages, you can better use the greater gas mileage afforded by hydrogen.
Note, however, that you could theoretically build a hydrogen engine as large as the F-1A in terms of thrust. The RS-68 (also a gas generator cycle engine) on the Delta IV vehicle puts out around three quarters of a million pounds-force thrust so that's pretty big. Also, back in the 1960's, there was conceptual design work performed on an enormous hydrogen fuel, gas generator cycle engine called the M-1. On paper, that behemoth put out 1.5 million pounds-force of thrust just like the F-1 on the Saturn V. But that project was abandoned and here's why: hydrogen is very, very light so if you want to carry any appreciable amount, you need to have truly huge tanks. Huge tanks mean huge stages. Huge means heavy. Eventually it becomes a game of diminishing returns at the vehicle level.
What this discussion of J-2X and F-1A (and RS-68 and even M-1) shows you is the extreme versatility of the gas generator cycle. It can be used with nearly any reasonable propellant combination and it can be scaled from pretty darn small to absolutely enormous.
Shaving with Occam's Razor
Occam's Razor is the notion that one should proceed with simplicity until greater complexity is necessary. Along these lines, I will introduce you to a simpler engine cycle: the expander cycle. For this engine cycle, you do not use a gas generator to drive your turbine(s) so you don't have a second, separate combustion zone apart from the main combustion chamber. That makes everything simpler. Instead, you use only the heat gathered in the cooling the thrust chamber assembly (i.e., the main combustion chamber walls and that portion of the nozzle regeneratively cooled). See the schematic below.
See? I got rid of not just the gas generator but also the two valves that fed the gas generator. That's huge in terms of simplification. And whenever you can make an engine simpler you’ve usually made it cheaper and more reliable just because you have fewer things to build and fewer things that could break. Cool!
Here, however, is the problem: How much power do you really have just from the fluid cooling the walls? The answer can be found by looking at the table and seeing, for example, the RL10 thrust output is less than one-tenth of J-2X. You just can't pull that much energy through the walls. There have been attempts to increase heat transfer by various means including making the main combustion chamber longer than typical so that you have more heat transfer area or even by adding nubs or ridges onto the wall to gather up more heat. Using the longer chamber notion, the European Space Agency is working on an engine called the Vinci that almost doubles the thrust output from the RL10, but getting much further beyond that is darn tough. Also note that hydrogen is a wonderful coolant based upon its thermodynamic properties. Being a wonderful coolant means that it picks up a lot of heat. It is difficult to imagine using the expander cycle engine with another fuel beside hydrogen (though maybe methane might work … haven't examined it).
On the plus side, in addition to the simplicity, what the cycle shown offers is what is called a "closed cycle" meaning that no propellants are thrown overboard other than through the main injector. In a gas generator cycle engine, after the gas generator combustion gases pass through the turbine(s), it's dumped into the nozzle (or, in other schemes, dumped overboard in other ways). Any propellants or combustion products that do not exit the rocket engine through the main injector and through the main combustion chamber throat represent an intrinsic loss in performance. "But," you'll say, "the specific impulse for the RL10 and the J-2X in the table are the same." Well, that's a little bit of apples and oranges because it's based upon the nozzle expansion ratio. Another model of the RL10, the B-2, has a much larger nozzle extension and the vacuum specific impulse for that model is over 462 seconds (minimum). The European Vinci engine that I mentioned above has a projected vacuum specific impulse of about 465 seconds. Those are darn impressive numbers that make the mouths of in-space stage and mission designers drool.
A couple of final notes about the expander cycle engine. First, the RL10 is not quite like the schematic shown. It only has one turbine with one pump driven directly and the other pump driven through a gear box. Thus, the OTBV goes away (making it even simpler!). Second, there are versions of the expander cycle engine concept that are not closed cycles. In these versions, you dump the turbine drive gas overboard in a manner similar to what you do in a gas generator cycle. You are still using the heat from the chamber walls to drive the turbine(s), so it's still an expander, but with an overboard dump you can also leverage a larger pressure ratio across the turbine(s) and thereby get a bit more oomph out of the cycle. You sacrifice a bit of performance for more oomph. The Japanese LE-5B engine is an open expander cycle engine like this (also called an "expander bleed" cycle).
"We do these things not because they are easy…"
So, you've seen the incredibly versatile gas generator cycle engine. And, you've seen the simple yet limited expander cycle engine. So what do you do if you say, "The heck with it, I want the Corvette"? What if you want a closed cycle, high performance engine not limited to lower thrust levels and you're willing to accept consequent greater complexity? The answer is staged combustion. Below is a simplistic schematic for a staged-combustion engine.
CCV = Coolant-Control Valve
PBOV = Preburner Oxidizer Valve
In a staged combustion cycle engine, we rename the gas generator and call it the "preburner." The biggest difference between a gas generator cycle and a staged combustion cycle is what you do with the turbine exhaust gases. In a gas generator cycle, the turbine exhaust gases effectively get dumped overboard. In a staged combustion cycle, the turbine exhaust gases get fed back into the main injector and get "burned again." This is possible since the combustion in the preburner is off from stoichiometric conditions, meaning that in addition to combustion products you also have lots of leftover propellant (either fuel or oxidizer depending on the scheme). The leftover propellants from the turbine exhaust then become part of the mix of propellants in the main combustion chamber.
That sounds simple, right? It's just a twist on the gas generator cycle theme, right? Well, there are larger implications. First, think about the pressure drops through the system. On a gas generator cycle engine, the pressure in the gas generator can be lower than the main chamber. After all, the downstream side of the turbine(s) is effectively ambient, external conditions. In a staged combustion cycle, the preburner pressure has to be substantially higher than the main chamber pressure sitting downstream of the turbine(s) or you don't get enough flow to power the turbine(s). Insufficient turbine power and the cycle doesn't work. So, in general, a staged-combustion cycle engine has higher system pressures than a gas-generator cycle engine of comparable size. Next, think about starting the system. In a gas generator cycle engine, the two combustion zones are effectively disconnected. In a staged combustion cycle engine, the two combustion zones are on either side of the turbine(s) so there is effectively communication between these two zones. Now, try to imagine getting these two combustion zones ignited and up to pressure and the turbine(s) spun up to speed in an orchestrated manner during the start sequence. It ain't easy.
So, what do you get for this complexity and higher operating conditions? Well, you get a closed cycle, high performance, and high thrust engine design choice. The RS-25 (SSME) is the American example of such an engine. If you put a higher expansion ratio nozzle on the RS-25, just as with the RL10 discussion, the specific impulse value would be as much as ten seconds higher than J-2X. However, if you go out and find a schematic of an SSME, what you'll see is a heck of a lot more complexity than even I've shown in my simplified sketch. Because the pressures are so high, there are actually four separate turbopumps and a boost pump in the SSME. The design relies on putting pumps in series to achieve the necessary pressures and fluid flow rates through system. And, the SSME has not one but two separate preburners, one for the high pressure fuel turbopump and one for the high pressure oxidizer turbopump. It's a very complex engine, but it has extraordinary capabilities.
The RS-25 (SSME) is a staged combustion cycle engine with hydrogen as the fuel. The preburners are run fuel-rich such that the generated gases contain excess hydrogen for injection in the main chamber. Back in the days of the Soviet Union, they developed a whole series of staged combustion cycle engines that instead used kerosene as the fuel. In these engines, the preburner is run oxidizer-rich so that the gases run through the turbines and then through the main injector have excess oxidizer to be used for final combustion in the chamber. The Russian-supplied RD-180 that is currently used for the Atlas V launch vehicle is an example of such an engine. It too is an extremely complex, high pressure, and high performance engine.
So, staged combustion cycle engines are not easy. Their complexity and operating conditions suggest, generically, greater expense and lower reliability. But if you can make the trade-off between high performance and the adverse issues, then they can function quite impressively. Nearly thirty years of Space Shuttle flights are an indisputable demonstration of this fact.
Just One Bolt
Can you imagine opening a hardware store and selling just one kind of bolt? That would be it. One brand. One diameter. One length. And just one bin full of identical versions of this one bolt in your store. It sounds really kind of stupid. The unavoidable truth is that you need different bolts for different applications. It's kind of like trying to imagine telling the Olympic gymnastics team that they now had to play basketball and the basketball players to do gymnastics. I don’t know about you, but I'd love to see Lebron James have a go at the pommel horse.
Well, over the last fifty-plus years, we've developed different rocket engines and rocket engine concepts for a variety of different applications. Just one design does not fit all applications. Each design has advantages and disadvantages. If you can understand the basics of what I've discussed in this article, however, then you will have a fundamental understanding of at least 90% of the engines spanning that fifty-plus years of history. And that, in turn, might help you better appreciate why one bolt is chosen over another or why, for example, shot-putters tend to be a bit more beefy than cyclists.
J-2X Progress: Once Upon a Time at Stennis...
Posted on Jul 09, 2012 03:06:23 PM | William D. Greene
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…"]
J-2X Extra: Human-Rated Chili
Posted on Jun 26, 2012 03:59:01 PM | William D. Greene
I enjoy cooking. Most people think that when I say that, it's because I'm an engineer by training, that I like cooking for the structured notion of a recipe and exactly measuring things out and the chemical precision of mixing that with this, at this speed, under these conditions, with these implements, and then forming it all together with a specified heat input over a given time using appropriately sized and shaped pots and pans optimized for uniform heat transfer, blah, blah, blah, blah…
But, that's way, way off from the truth.
Actually, I like to cook things that allow for, let's say, "significant organic creativity." I make a mean vegetarian chili, but you can be sure that it will be different every single time that I make it since it's always from memory and my memory ain't what it used to be. I wing it. And that's fun. And even though it's fun and even though the details vary slightly, it's been good every time (so far). The worst side effect that I could attribute about any particular version might be a bit of heartburn (properly mellowing and blending habanero peppers is an imprecise art form I have not yet consistently mastered).
So, what does my free-form chili cooking this have to do with J-2X? Believe it or not, I want to talk about one of the adjectives that we frequently apply to the J-2X engine: "human-rated." What does that mean? We use that term (or the older, less politically-correct formerly used term "man-rated") all of the time and, for the most part, those of us within our little clique understand the general context of its meaning. But if you asked any of us to explain, you'd likely get a wide variety of different, complex, and mostly correct yet often partial answers. I am no genius and, despite all odds, I will do my best to provide a reasonably complete framework for a definition so as to help you better understand the J-2X engine.
And, it will come back to my cooking analogy. Really.
First, we need to recognize that there is really no such thing as a "human-rated rocket engine." That is shorthand terminology that ought to be written out as: "a rocket engine that could be suitable as part of an overall, human-rated launch system." Think of it this way: Let's say that you had a total junker of a car but you installed one perfectly pristine, top-quality piston. Do you now have a good car or do you still have a junker? You'd still have a junker, of course. Or, let's say that you had a really nice car but all of the spark plugs were corroded, eroded, and barely functional. Do you still have a nice car? Well, maybe the paint job is pretty and the stereo sound is clear, but it's not going to get anywhere quickly, reliably, or efficiently with bad plugs. The point is that no single element of something as familiar as an automobile makes it complete and good and, in an analogous manner, no single element of something as large as a launch architecture is, in itself, human rated. The whole system is rated for human spaceflight because the system as a whole, as well as its constituents such as the J-2X, meet certain standards and processes that we'll discuss below. We call the J-2X "human-rated" as a shorthand way of saying that it could be part of a human rated architecture consisting of the rest of the vehicle, ground operations, mission control, and exceptionally well trained ground and flight crews, etc.
Second, let's think about the adjective term "human-rated" itself and its definition. What does that mean? It means simply this: the estimated risk is acceptably low so that we can responsibly decide to put human beings into the vehicle for launch. Again, we can relate this to automobiles. When you drove to work today, you took a risk. Unfortunately, auto accidents happen on the roads and highways and, more unfortunately, despite all of the protective apparatus built into our cars, people do sometimes get hurt in these accidents, or worse. But you accepted that risk and drove to work anyway. You judged your auto to be sufficiently safe. You judged that the roads were well paved and properly marked, that the police were properly monitoring bad and endangering behavior on the roads, and that the weather was clear enough to allow for safe operation of your vehicle. Thus, your "drive-to-work system" was, today, according to your judgment, "human-rated" for you. You weighed the risks -- consciously or subconsciously -- and decided to accept these risks and make the trip.
Spaceflight is ten thousand times more complex than driving to work, but the rationale is entirely analogous. The "fly-to-space system" (note again it's a "system" not just a vehicle) is "human-rated" when we judge the risk to be acceptable in light of the potential rewards. The important and fundamental point is that, in the end, it is a judgment. Sometimes, for example, we accept more risk because we judge that the potential rewards are that much more significant. Think back to the early days of human spaceflight. I can guarantee that there is no way in heck that we would today put an astronaut into some of those early vehicles. We would not today consider those early systems to be human-rated by our current standards. But at that time, we as a nation accepted the risk and, by the way, achieved extraordinary milestones. Today, our objectives and potential rewards are different and so our judgments with regards to risk are accordingly different.
So, if it's all just a matter of judgment, then doesn't that mean that there really is no such thing as "human-rated"? No, I would strongly disagree.
Here is where I get back to my cooking analogy. While my chili may have slightly different constituents each time that it's made, and while it might taste a bit different each time, there is no question as to whether it is chili. I use my expert cooking judgment to combine the essential ingredients into a recognizable and tasty product (with or without subsequent heartburn). When we talk about an engine being "human-rated," we too are not basing that judgment upon a fixed recipe. We are basing it upon a combination of essential ingredients and expert judgment.
If you're wondering whether NASA maintains some kind of formal recipe for human rating, I refer you to NASA Procedural Requirements (NPR) 8705.2, revision B (effective May 2008), "Human-Rating Requirements for Space Systems." While this document is helpful, in a general sense, with regards to what technical and programmatic areas to consider, it is written at a very high level, i.e., at the "fly-to-space system" level. As such, it does not offer a great deal of rocket-engine-specific information. This, in my opinion, is exactly as it should be. The actual making of the chili should be left to the expert cooks. Even NPR 8705.2 makes it quite clear that the intent of the document is only to establish a framework within which "human rating" takes place. It is not intended to be a step-by-step recipe book for the many, many diverse parts of a human spaceflight system.
What then are the essential ingredients for a human-rated engine? Not surprisingly, the answer can be thought of as somewhat following the life cycle of an engine development project.
Design and Development
Specific technical requirements -- There is a small handful of specific technical requirements that effectively flow down from NPR 8705.2B and impact the engine design. One is the requirement that, where appropriate and where it can be shown to increase reliability and safety, we should use redundant systems. On the J-2X, the clearest manifestation of this is the use of an engine controller with two channels. Should one channel fail (as even heavy-duty computer systems sometimes can), the other channel can take over and continue safe operation. Another specific requirement at the system level is that there exist abort systems that allow the crew to escape from a bad situation on the vehicle. This requirement decomposes to a requirement on the J-2X for a redline health monitoring system that shuts down the engine in the event of an imminent failure and notifies the vehicle of this shutdown. This thereby allows the crew the opportunity to perform an abort.
Design, construction, workmanship standards -- Not surprisingly, we don't start from scratch every time that we sit down to design something. We know how to do things. We have lessons learned. We have rules of thumb. And, at the top of the list, we have standards. These are specialized requirements documents that focus on specific, narrow technical areas. For example, NASA-STD-5012 tells you what you should do for the structural design of a rocket engine. It lays out the essential analyses to perform, the way that the environments should be evaluated, and what factors of safety are appropriate. For J-2X, we had over thirty different standards that were (and are) part of the requirements imposed upon the engine design details, design processes, fabrication processes, and testing scope and procedures.
Even here, however, after you impose a standard you have to acknowledge the fact that there can exist more than one way to do things and do them safely. For example, on J-2X we imposed a structural design standard that, at a lower level, imposed a standard for how fasteners (i.e., bolts and nuts) are properly lubed and torqued. In order to investigate this issue, we set up a mini-test program to better understand the results from the different methods. It kind of sounds silly, but fastener torque is extremely important in high-pressure systems and proving that the contractor process was equivalent and safe could save us money in the long run since it is a standard procedure for them. So, we had a guy follow the procedures several times and we measured the strain induced into a series of bolts by the applied torquing method. The measured strain was converted to applied force and this thereby validated the procedure. Across the spectrum, we had a number of similar examples where we interpreted the technical intent and purpose of a detailed requirement and, working with our contractor, found the best way to comply.
System safety program -- As an engineer, the question foremost in your mind is always, "How can I make this thing work?" Without that mindset, we would never get anywhere. However, when dealing with something as complex and as potentially dangerous spaceflight, you must go beyond this level of thinking and must also continuously ask yourself, "What could go wrong with this thing and how do I mitigate that potential as much as possible?" In the most basic sense, this is the motivation for developing a system safety program. As part of the engine design and development process, you look at this issue from two directions.
First, you look at the piece-part level and ask, "What could break, how or why, and what would be the effects?" That's a reliability analysis. You look at all of the pieces and figure out what circumstances could result in something not working as intended. Could the design be mistaken because we didn’t understand the loads? Could the loads go off nominal because of some unusual flight situation? Could the manufacturing of that piece go awry so that you don't have the intended design margins in the actual, physical part? And, for all of these questions, you have to provide answers as to how best to ensure that the part won’t actually break during operation.
Second, you start from the other end. You start with the grim notion that you've failed and that the crew didn't make it. From there you work backwards and figure out how and why that situation could take place. This process grows into a tree of circumstances and possibilities and is called a hazards analysis. Was it an explosion? If so, where did the fuel and oxidizer and ignition source come from? If the fuel came from tank, then how did it escape? Was it instead something having to do with navigation? Or maybe there was a weather-related issue, perhaps, say, lightning?
Obviously, in many places these two assessments eventually meet in the middle. The one starts at the bottom and works upwards. The other starts at the top and work downwards. When they meet, then you know where throughout your system are your critical points. In some cases this drives design features, special inspection requirements, or, for example, in the case of lightning protection, the design and construction of a launch pad system for dealing with the hazard. This overall effort allows you to prioritize your efforts to ensure safety and, in the operational phase, potentially apply greater attention prior to committing to launch.
Test and Evaluation
Structured verification planning and reporting -- Believe it or not, we don't march into an engine test program all willy-nilly and make a bunch of smoke and fire just for the sake of impressing our friends. We do it to generate and collect data. The data that we collect largely goes towards the systems engineering endeavor known as requirements verification. Verification is defined as the process of demonstrating that the product design -- in our case an engine -- is in compliance with imposed requirements. Verification can, and does, take a number of forms. Testing is one form. Analysis and inspection are others.
Note that the "structured" part of the "structured verification" title above is a key consideration. You must lay out plans saying, "Here is my requirement and here is what I plan to do to prove that I meet it." Then, based upon peer review of experts, this plan can be approved or modified. This is an essential part of the whole judgment aspect of human rating. If I demonstrate that I meet the requirement with one engine on one test, is that good enough? If not, how many engines or tests do I need? Or, if it's verification by analysis, do you agree with the analysis methodology that we propose to use? Do you concur with the assumptions and the simplifications inherent in any analysis method? The whole process, when properly approached, has the flavor of the classic scientific method. The hypothesis is that the product meets the requirement and then you set out to prove that hypothesis.
Smart people with backgrounds in mathematics inevitably jump into the conversation here and declare the supremacy of statistics. Using statistical analysis, we can determine how many samples and tests are necessary to achieve a mean and variability assessment at a given confidence level. Unfortunately, as good as those methods might be, we can never come close to affording the kinds of programs that a purely statistically based assessment would suggest. Maybe back in the day we could afford to build and test 100 engines before we're ready to fly, but today our constraints are to accomplish the same level of risk mitigation with an order of magnitude fewer samples. We have to be wiser and more efficient, and yet still have sufficient confidence to declare that the design meets its requirements.
Test, test, test, and then test some more -- Now, after having discussed a fundamental motivation for testing engines, i.e., requirements verification, you have to get down to the nuts and bolts of the issue. You must test and you must do it a lot. Yes, "a lot" is not what you'd call a scientific term, but it can be decomposed. "A lot" means that you cover your verification plans in terms of samples and repeat examples. It means that you push things beyond normal operation to prove margins. You test longer -- both single run and cumulative on a given engine, both starts and seconds -- than any flight engine could possibly ever see. And throughout this process, you continuously learn things that you didn't know that you didn't know. While it is theoretically possible that we could design an engine, put it into test, and find that we'd properly characterized every environment and every engine response to those environments, but I've never seen such a case and nobody that I know have ever heard of such a thing. Engine testing is always an education.
The other aspect of testing that is sometimes categorized separately is teardown and detailed inspection of the hardware afterwards. If you predicted that something wasn't going to crack and, upon teardown, you find a crack where it shouldn't be, then you're not as smart as you thought you were (a phrase I've used before). If you tear down and find that something was rubbing in a valve or a turbopump, then that might be an issue. Or, instead, it might have been planned that way. You look for discoloration that might suggest unexpected operational conditions or potential changes in material properties. You check dimensions of everything to make sure that you didn’t deform pieces or possibly lose material that was consumed by the engine. Thus, while you collect lots and lots of data during the engine tests, it is also the data that you collect after the testing is complete that contributes substantially to your understanding of the design and its safe operation.
Quality processes -- Twenty-some years ago, the Ford Motor Company had a motto that they used in advertising: "Quality is Job One." With all due respect to that venerable motor company, those of us in the rocket world have known this for a long, long time.
When we certify an engine design and say that it is "human-rated," that is a contingent description. It is contingent upon future flight engines being produced in the same manner and to the same detailed workmanship standards as the design that you certified. That means that the fabrication and testing processes are the same, the materials are the same, the people doing the work on the pieces have had the appropriate training, and that the finished parts have been scrutinized to the same inspections and inspection standards. And, if things can’t be exactly the same (for example, vendors can change over time), then you must have a process in place to assure equivalence between what you had before and what you're going to use new.
Also, should something go awry during the manufacturing or assembly of any part -- and things always go awry to some degree at some point -- you need to have processes in place to identify what went wrong, how to avoid that issue in the future, and what to do with any hardware that was exposed to the issue. Can you fix it and still meet your requirements and drawing specifications? Or, do you have to scrap the part because it can't be saved?
These considerations are all part of a good, solid quality system.
Configuration management -- The first cousin of quality assurance is configuration management. While it sounds like a simple premise, this discipline deals with making sure that the exact, particular pieces on the vehicle are the exact, particular pieces that you intended to put on the vehicle. This means, for example, that every bolt on the engine is suitable for a flight engine. No, not every bolt has a serialized part number, but they are segregated by lots. Lots intended for flight usage are subjected to a stringent quality processes and must, therefore, be kept separate from any similar-looking bolts that might not meet the high standards for flight. Plus, of course, we track throughout their lives the history of our serialized assemblies like turbopumps, combustion chambers, nozzles, ducts, lines, controllers, valves, etc., along with their associated documentation. And engine is composed of thousands of parts and, one way or another, we track them all.
The combination of a good quality assurance system and a good configuration management system guarantees that what you have delivered and put on the launch vehicle is exactly what it is advertised and intended (and needs) to be.
That's it. Those are, in my opinion, the key ingredients for human rating.
So, getting back to cooking. In order to make vegetarian chili, you need tomatoes, beans, and chili powder. That's it. But chili made with just these ingredients would be terrible. I add peppers (of multiple varieties) and onions and garlic and other spices. Corn can add a nice sweetness. Sometimes I sauté chopped portabella mushrooms and toss them in. Beyond that, I've been known to add all kinds of oddball stuff including, once, green beans. And, in the end, it's good. I promise. That's because I've made it probably thirty or forty times over the years and therefore I am a subject matter expert (within my tiny culinary world). Solid, well-defined ingredients and expert judgment inform my chili.
In order to have a "human-rated" rocket engine, all of the topics that I mention above represent the key, essential ingredients: (1) a few, specific human-rating design requirements, (2) a set of established design, construction, and workmanship standards, (3) a thorough safety program, (4) a structured verification process, (5) system testing campaign, (6) a solid quality assurance system, and (7) a reliable configuration management system. They are all necessary. And certain bounds, limits, or standards can be established (and are documented) for all these various disciplines and undertakings, but an exact, repeatable, or universal, step-by-step recipe is extremely difficult to conjure up. Just like my chili, the details of how, when, and why an engine is "human rated" fall within purview having good key ingredients and then applying expert judgment.