SSME – Liquid Rocket Engines (J-2X, RS-25, general)

J-2X Progress: Shaking up the Night

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

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

Then, the engine started.

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

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

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

Liquid Engines Extra — Introducing LEO

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

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

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

…wait for it…

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

So, let me introduce you to LEO —

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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!