Tag Archives: space launch system

LEO Progress: RS-25 Adaptation

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

Liquid Engines Extra — Introducing LEO

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

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

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

…wait for it…

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

So, let me introduce you to LEO —


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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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


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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

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

…oops, wrong movie. 

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