Firing Up the Shuttle: 25 Years Since the Last Flight Readiness Firing (Part 2)

Each Flight Readiness Firing (FRF) presented the opportunity to extensively test the Space Shuttle Main Engines (SSMEs) in as close to real-flight operations. Photo Credit: NASA

A quarter-century ago, on 6 April 1992, the roar of three Space Shuttle Main Engines (SSMEs) echoed across the marshy landscape of the Kennedy Space Center (KSC) in Florida, as Endeavour showcased her muscle, ahead of her maiden voyage into orbit. Built as a replacement for her lost sister, Challenger, the new vehicle would go on to fly 25 missions—supporting the first (and only) three-person spacewalk, servicing the Hubble Space Telescope (HST) and building the International Space Station (ISS)—and cement her credentials as the fourth-most-flown member of the shuttle fleet. Within 22 seconds of Main Engine Start, her engines fell silent, as planned, to close out the final Flight Readiness Firing (FRF).

As outlined in yesterday’s AmericaSpace history article, FRFs were executed before each orbiter’s maiden voyage, and were performed twice by shuttles Challenger and Discovery. The intention was to impose approximate launch-like conditions on the SSMEs, as well as testing the Auxiliary Power Units (APUs) in high-speed mode. Visually and acoustically, it was perhaps the closest example—other than a launch itself—of the orbiters straining against their shackles, yearning to fly. All control elements of the Main Propulsion System (MPS) were required to hold pressure in the engines and the External Tank (ET) and the flight control instrumentation was expected to provide proper throttling and gimbaling functions. This would serve to validate the integrated performance of the Space Shuttle “stack” and the compatibility of the on-board General Purpose Computers (GPCs) with ground-based computer systems.

Endeavour’s FRF on the morning of 6 April 1992 proceeded relatively smoothly, with the notable exception of high vibration levels in one of the SSMEs’ high-pressure liquid oxygen turbopumps. Taken as an indicator of hydrogen ingestion into the fuel injector, it was decided to replace the engines with another set.

Challenger’s three main engines ignite on 18 December 1982 for the first Flight Readiness Firing (FRF) of STS-6. Photo Credit: NASA

It was the final FRF of the shuttle era, coming 11 years after the engines of Endeavour’s sister ship, Columbia, roared to life on 20 February 1981. Interestingly, the second orbiter to undertake an FRF actually did so twice. At 11 a.m. EST on 18 December 1982, shuttle Challenger was almost ready for her maiden voyage, with launch targeted to occur in the second half of January 1983. With a minute to go before Main Engine Start, the Public Affairs Office (PAO) announcer reeled off the timeline, as Challenger came alive and her vehicle systems were brought online.

“T-1 minute and counting…the firing system that releases the sound suppression water onto the pad has been armed…T-50 seconds and counting…T-45 seconds and counting…T-40 seconds and counting; SRB development flight recorders are being turned on…T-37…gaseous oxygen vent arm will not be retracted on this particular test…T-31 seconds, we have a Go from LPS [Launch Processing System] for auto-sequence start…[Challenger’s] four primary flight computers taking over control of the terminal count…final LPS command for engine start will occur at approximately 10 seconds…T-15 seconds and counting…”

At this stage, the relative silence on the pad changed markedly. Firstly, the sound suppression system gushed water across the launch pad. “T-10…Go for Main Engine Start…we have Main Engine Start…” as the now-familiar sheet of orange flame gave way to a trio of shock diamonds from the three SSMEs, combined with a thunderous roar and vast cloud of smoke. The engines ignited in a ripple-like sequence, starting up at 120-millisecond intervals, reaching 90 percent of rated performance within three seconds and hitting 100-percent at zero.

“T-0, engines throttled at 100 percent, all engines up and burning,” came the second-by-second updates from PAO. “T+5 seconds, engines continuing to burn…T+10 seconds…twelve…first [engine] cutoff at T+15 seconds…[Number One] engine cutoff…and engines Two and Three also cutoff at 16.8 seconds…T+25 seconds; GLS safing now in progress…” However, the FRF was not yet over, for the APUs were run up to T+2 minutes in high-speed mode, after which GLS safing of the vehicle was completed.

According to NASA Launch & Landing Operations Director Al O’Hara, it was anticipated that about 48 hours after the test, by 20 December 1982, the initial data was expected to be in place, after which the actual physical inspection of the SSMEs could commence. In the immediate aftermath of the FRF, O’Hara described it as “a resounding success”, but it later became clear that everything did not run according to plan. “Let me caution you that this is based upon the real-time information from the firing room and from the support rooms that we got on-net about 30 minutes after T-0,” he told journalists later that morning. “So as the day progresses and more information get available, that may change.”

And indeed it did.

During the test, engineers detected levels of gaseous hydrogen in the shuttle’s aft compartment which grossly exceeded allowable limits. When it proved impossible to pinpoint the cause or location of the leak, the Mission Management Team (MMT) elected to perform a second FRF. New instrumentation was installed inside and outside Challenger’s aft fuselage to better determine if the leakage was from an internal or external source, with suspicion initially focusing on the latter possibility, because vibration and current had found their way behind the SSMEs’ heat shields.

Extra sensors and a higher-than-ambient pressurization level were installed to prohibit penetration by “external” hydrogen sources and the second FRF, lasting 23 seconds, took place on 25 January 1983. It too revealed high concentrations of hydrogen gas, necessitating the replacement of one of the SSMEs and, subsequently, the replacement of the other two engines, due to detection of cracked welds, fractured fuel lines and generic “seepage” in an inconel-625 tube within the ignition system. Described by NASA Associate Administrator for Space Flight James Abrahamson as “a real detective job”, a third FRF was briefly considered, but by mid-February this proved unnecessary.

Over the course of the next three years, two more orbiters—Discovery and Atlantis—also embarked on their FRF rite-of-passage, ahead of their maiden voyages. Discovery, which wound up as NASA’s most-flown shuttle, saw her SSMEs burned to full power on 2 June 1984. Operationally, her FRF was not dissimilar to its predecessors, with the exception that the forward motion of the vehicle (nicknamed “the twang”) was more pronounced, although well within limits. And on the morning of 12 September 1985, Atlantis’ engines roared for her own FRF, ahead of her inaugural mission a month later.

All five orbiters performed FRFs in the weeks preceding their respective maiden voyages, with Challenger and Discovery performing two apiece during their careers. Photo Credit: NASA

It was expected that this would be the last, first-time FRF for a member of the shuttle fleet. However, with NASA and the U.S. Air Force planning to initiate Space Launch Complex (SLC)-6 at Vandenberg Air Force Base, Calif., for West Coast-based flights from the summer of 1986, it was intended that Discovery would perform another FRF prior to Mission 62A. In the aftermath of the Challenger accident in January 1986, Vandenberg missions were canceled, but Discovery was tasked with her second FRF in August 1988, prior to STS-26. This met with several delays, including a false start on 4 August, when a sluggish SSME valve led to an abort at T-6.6 seconds. Some consternation was apparent in the voice of the PAO, for the engines’ hydrogen burn igniters had already kicked into action, but no Main Engine Start occurred.

Six days later, on 10 August, there were no such problems and the shuttle’s engines fired smoothly for 20 seconds of continuous thrust. This allowed NASA to clear another milestone in getting the fleet back to post-Challenger operations in late 1988. The principal reason for an FRF before STS-26 was to evaluate more than 200 SSME and other vehicle modifications and, with Endeavour’s test in April 1992, no further plans were laid. Significantly, STS-114—which marked the post-Columbia return-to-flight mission in July 2005—did not require an FRF, because the External Tank (ET) had been the principal focus of modifications, rather than the SSMEs themselves.

“On every FRF that we conducted, we learned something new about the vehicle, which made our process and flight hardware better,” said deputy shuttle processing chief engineer Jorge Rivera. “It’s definitely a good practice in reducing the risk of the actual flight.” Critically, Rivera added that the FRF was in keeping with the “test as you fly, fly as you test” mentality. “Test is the best control or mitigation for hazardous conditions that could impact the mission,” he explained. “Subsystems that tested fine in isolation may interface with each other in a different way, which could create a bigger problem.”

Across their 30-year-plus career in the Space Shuttle Program, the SSMEs posed major headaches for NASA. Agonizing failures and explosions in the late 1970s, coupled with no fewer than five Redundant Set Launch Sequencer (RSLS) abort situations—premature engine shutdowns, with crews aboard, seconds from T-0—contributed to a pre-Challenger belief that they would be the lead contender in any launch failure. However, the SSMEs were extensively modified and proved themselves to be one of the safest and most reliable elements of the shuttle system. Their contribution continues, for heavily upgraded and reconditioned SSMEs will soon form the Main Propulsion System (MPS) of the Space Launch System (SLS), when it flies its inaugural voyage in the second half of the present decade.

 

This is part of a series of history articles, which will appear each weekend, barring any major news stories. Next week’s article will focus on the 15th anniversary of STS-110, a shuttle mission which saw Jerry Ross become the first human to launch into space, from Earth, a total of seven times.

 

 

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11 comments to Firing Up the Shuttle: 25 Years Since the Last Flight Readiness Firing (Part 2)

  • James

    “However, the SSMEs were extensively modified and proved themselves to be one of the safest and most reliable elements of the shuttle system. Their contribution continues, for heavily upgraded and reconditioned SSMEs will soon form the Main Propulsion System (MPS) of the Space Launch System (SLS), when it flies its inaugural voyage in the second half of the present decade.”

    Thank you Ben Evans! Your two historical articles on the “heavily upgraded and reconditioned SSMEs” and “Flight Readiness Firing” are interesting and provide some food for thought.

    Eventually, a future extensively modified and evolved Aerojet Rocketdyne RS-25 engine that is capable of being re-started in space many times might turn the core of the SLS into a valuable space tug or even a huge Lunar and Mars Lander.

    Note:

    “The Space Shuttle Main Engine, on the other hand, would have required extensive modifications to add an air-start capability and be able to restart in a near-vacuum. The near-vacuum restart capability was needed for the Ares I as it was intended to fly an Earth orbit rendezvous, and because the Orion spacecraft has limited fuel reserves. Due to these design issues, a modified Space Shuttle Main Engine would have to be ‘pre-fired’ in a manner similar to the ‘Main Engine tests’ conducted on the Space Shuttle Main Engines prior to the maiden flights of each NASA orbiter, including the STS-26 return to flight in 1988.[12]”

    From: ‘Ares I’ Wikipedia
    At: https://en.wikipedia.org/wiki/Ares_I#J-2X_engines

    Adding propellant in LEO, Lunar, and Mars orbit to the SLS’s core’s tanks would need to be done, but the potential opportunities offered by such evolved RS-25 engines with vacuum restart capability powering an appropriately modified core of the SLS do seem awesome.

    And of course with possible future reusable SLS boosters landed back on Earth, those modifications could turn the SLS into a fully reusable launch system for Lunar and beyond Cislunar missions and the contribution of the modified versions of the SLS could continue for decades.

    Note:

    “The Soyuz launcher was introduced in 1966, deriving from the Vostok launcher, which in turn was based on the 8K74 or R-7a intercontinental ballistic missile.”

    From: ‘Soyuz (rocket family)’ at Wikipedia

    “The R-7 family of rockets (Russian: Р-7) is a series of rockets, derived from the Soviet R-7 Semyorka, the world’s first ICBM. More R-7 rockets have been launched than any other family of large rockets.”

    From: ‘R-7 (rocket family)’ Wikipedia,

    Yep, the R-7 first flew on May 15, 1957 and the latest version in the family is the Soyuz-2 that probably has many decades of launches in its future.

    Maybe modified versions of the SLS’s core powered by evolved Aerojet Rocketdyne RS-25 engines and new reusable boosters might also be launched for a very long time into the future.

    Time will tell.

  • This time James’ vapid comment was somewhat useful. Useful because it pointed out a glaring error in the Wikipedia Ares I entry. This ties in with James’ notion of using the SLS core as some sort of interplanetary um -thing.
    I will edit the Wikipedia entry asap.

    PART I:

    …add an air-start capability and be able to restart in a near-vacuum. The near-vacuum restart capability was needed for the Ares I as it was intended to fly an Earth orbit rendezvous, and because the Orion spacecraft has limited fuel reserves

    No. There was never any requirement for the (CLV) Ares I upper stage engine to “restart”. Ever.
    The Ares I ascent profile was similar the Shuttle profile, the SRB would land in the Atlantic, then the cryogenic liquid upper stage would inject the Orion into an elliptical orbit with perigee intersecting the upper atmosphere. After Orion separation, the upper stage would burn up over the Indian ocean, while the Orion main engine would circularize Orion’s orbit.

    …Orion spacecraft has limited fuel reserves
    The original Constellaion Program Orion using a modified AJ10 had plenty of fuel to circularize its orbit, then rendezvous with the Ares V Earth Departure Stage / Altair stack.
    The current Orion uses an ESA Service Module with a (literally) left over Shuttle OMS engine, however, the SLS launched Orion will be put in a circular low earth orbit by the SLS upper stage.

    PART II

    James’ absurd notion of a “…modified and evolved Aerojet Rocketdyne RS-25 engine that is capable of being re-started in space many times might turn the core of the SLS into a valuable space tug or even a huge Lunar and Mars Lander. is tied in with the erroneous Wikipedia Ares I entry.

    The RS-25 (SSME) can not be restarted in space. The staged combustion cycle RS-25 is started by “head pressure”, the hydraulic pressure of over 500 tons of propellants under one earth gravity. There is no way around this, period.

    An interesting aside: if you look at the renderings of the first official Constellation Program Ares I design with the RS-25 engine, you’ll see it has a “lattice” interstage, similar to the Russian’s Soyuz launcher. This is because (1): starting an RS-25 is a very very involved, hour long chill-down & purge procedure which vents considerable amounts of explosive gas and hydraulic fluids. The open lattice would reduce the gas buildup. (2): The smaller (than Shuttle) Ares I upper stage tanks, would provide less head pressure to start the RS-25s, so the decision was made to save money on RS-25 modifications by using the SRB thrust to “fake” the RS-25 Shuttle hydraulic head pressure. This version of Ares I would do a “fire in the hole” ignition of the RS-25 upper stage engine, by starting the RS-25 a split second before SRB burn-out, the acceleration from the SRB would increase the head pressure of the upper stage feed lines. (3) the open lattice would reduce the back-pressue shock to the RS-25 nozzle, hopefully reducing more expensive mods.

    Needless to say, this “fire in the hole” air-start RS-25 had a very high “pucker factor” (in engineer parlance) and would require lots and lots and lots of actual flight testing to certify it for human use. For instance, the classic “pogo” oscillations from the big SRB would make the head pressure variable, requiring an elaborate damper and accumulator system. Bummer.

    When the price tag became apparent, going to the actual air-start designed J2 engine looked really good, they just needed a 5 segment SRB to get the whole thing to work. Also, the J2 was used on the Constellation Ares V upper stage, so “comminality” would be nice.

    Naturally, making a “cheaper simplified” J2-X work on the Ares I turned out to be way harder and more expensive than planned, so after years of government mega dollars burned, Obama was elected and Ares I with Constellation became history.

    PART III

    The James/Church notion of “…future reusable SLS boosters and…core of the SLS into a valuable space tug or even a huge Lunar and Mars Lander is beyond preposterous.

    (1) The SLS core does not go into orbit, it re-enters and burns up over the Indian Ocean.
    (2) If by some miracle you paid to reduce payload in order to put the core into orbit, you’d have –nothing useful.
    (3) As previously, the RS-25 can not be restarted on orbit.
    (4) If by some miracle you had restartable engines, the huge 84-ton core has no ullage motors. Ullage motors for this monster would weigh as much as most payloads anyway.
    (5) Along with ullage motors goes, attitude control and articulation, power, communication & navigation, thermal control & heat rejection for the electronics. Once on orbit, the foam insulation would turn dark and begin to pop off from UV and solar thermal. The list goes on.
    (6) The SLS core would make terrable upper stage Lunar-Mars um -thing for the above reasons, plus it would have an awful trust-to-weight ratio because it’s the core of a big-ass lauch vehicle. The core must withstand the gigantic masses, acoustic loads, dynamic loads, aerodynamic loads while it serves as a strongback for the SRB/core stack and then becomes and upper stage booster. NONE of these qualities are good for a space born vehicle.
    (7) Reusable?? The SLS core reaches (just shy of) orbital velocity, which means to bring it back, you’d have to do a “once around” the globe, then reenter the atmosphere like any s/c. Fine. Except the SLS core would need a Massive heat shield of some sort, attitude control system, Massive landing gear, and an all propulsive landing using engines which can not be restarted. Ok so far.
    (8) Ok, so assuming the hideously expensive, behind schedule (literally 20 years in the making) uber low flight rate, SLS survives being embarrassed by SpaceX & Blue Origin, the US government is going to say: “will OK, so SLS is a joke, but now let’s spend a gazillion dollars and another 20 years to make the core into a fabulous interplanetary wet-workshop Lunar-Mars Lander monstrosity -thing.

    Time will tell.

  • James

    And yet you accept that the ITS will someday be reusable over vast distances and in diverse environments?

    OK.

    I accept that an appropriately modified hydrolox core of the SLS and its ever evolving RS-25 rocket engines have an excellent chance of being reusable over vast distance and diverse environments.

    “(1) The SLS core does not go into orbit, it re-enters and burns up over the Indian Ocean.”

    That is mainly because it is currently planned to do it that way and does indicate the current or inherent capabilities of an evolved SLS core.

    “(2) If by some miracle you paid to reduce payload in order to put the core into orbit, you’d have –nothing useful.”

    An opinion based on a quite limited perspective concerning what future boosters will be used for the SLS.

    “Estimates in 2012 indicated that the Pyrios booster could increase Block 2 low-Earth orbit payload to 150 t, 20 t more than the baseline.[50]”

    From: ‘Space Launch System’ Wikipedia

    Much larger and reusable boosters are possible.

    Reusability adds weight to all launchers with reductions to payloads that may range from 30% to 50%.

    “(3) As previously, the RS-25 can not be restarted on orbit.”

    Not currently. But methods to do precisely that have been considered in the past.

    “It would be expensive, time-consuming, and weight-intensive to convert the ground-started RS-25D to an air-started version for the Ares I second stage.”
    From: ‘Space Shuttle main engine’ Wikipedia

    As I noted previously, reusability “adds weight to all launchers with reductions to payloads that may range from 30% to 50%”.

    However, reusability may lower mission costs and is worth careful consideration.

    “(4) If by some miracle you had restartable engines, the huge 84-ton core has no ullage motors. Ullage motors for this monster would weigh as much as most payloads anyway.”

    No “miracle” is needed, just excellent engineers.

    RL10 ullage engines would be useful for such a modified and reusable SLS core.

    “(5) Along with ullage motors goes, attitude control and articulation, power, communication & navigation, thermal control & heat rejection for the electronics.”

    Wow! Whatever happened to rocket engineers? These are resolvable issues.

    “Once on orbit, the foam insulation would turn dark and begin to pop off from UV and solar thermal. The list goes on.”

    And second stages of various launchers, including the SLS’s Exploration Upper Stage, can’t resolve these issues?

    Don’t use the current foam insulation.

    Eventually use a carbon-fiber based SLS core. The folks designing the ITS seem to believe they will be able to solve these issues.

    And/Or:

    “MIT discovered that by taking many small flakes of graphene and fusing them together they could essentially create a mesh-like structure that, while porous, retained graphene’s amazing strength properties. They used 3D plastic models to test what kind of a structure would be the strongest under pressure, and then arranged the graphene in the same manner. The resulting material is only 5% as dense as steel, but an amazing 10 times stronger.”

    From: ‘MIT just invented one of the strongest, lightest materials known to man’
    By Mike Wehner 1/7/2017
    At: https://www.yahoo.com/tech/mit-just-invented-one-strongest-lightest-materials-known-221206439.html

    “(6) The SLS core would make terrable upper stage Lunar-Mars um -thing for the above reasons, plus it would have an awful trust-to-weight ratio because it’s the core of a big-ass lauch vehicle. The core must withstand the gigantic masses, acoustic loads, dynamic loads, aerodynamic loads while it serves as a strongback for the SRB/core stack and then becomes and upper stage booster. NONE of these qualities are good for a space born vehicle.”

    Isn’t the strongback for the propellant loaded SLS stack actually the SRBS, which do not go into space?

    “(7) Reusable?? The SLS core reaches (just shy of) orbital velocity, which means to bring it back, you’d have to do a ‘once around’ the globe, then reenter the atmosphere like any s/c. Fine. Except the SLS core would need a Massive heat shield of some sort, attitude control system, Massive landing gear, and an all propulsive landing using engines which can not be restarted. Ok so far.”

    An evolved RS-25 could be restarted.

    And, Lunar hydrolox propellant in LEO could resupply the SLS core with propellant which could be used to significantly reduce reentry thermal stress issues.

    If need be, buy a modified New Glen heat shield.

    If need be, buy some modified New Glen landing legs.

    In the near-term, buy hydrolox propellant hauled up to LEO by New Glen launchers.

    (8) Ok, so assuming the hideously expensive, behind schedule (literally 20 years in the making) uber low flight rate, SLS survives being embarrassed by SpaceX & Blue Origin, the US government is going to say: “will OK, so SLS is a joke, but now let’s spend a gazillion dollars and another 20 years to make the core into a fabulous interplanetary wet-workshop Lunar-Mars Lander monstrosity -thing.

    Nahh. The “Lunar-Mars Lander monstrosity –thing” is a core of the SLS and is based on hydrolox and is thus quite suitable for landing over 1,000,000 lbs of cargo on the Moon, Mars, or Ceres. Water everywhere means hydrolox everywhere.

    Hydrolox is a much more efficient propellant than methalox. Hydrolox got us to the Moon in 1969, not methalox.

    And oddly enough:

    “The Interplanetary spaceship will operate as a second-stage of the orbital launch vehicle on Earth-ascents—and will also be the interplanetary transport vehicle for both cargo and passengers— capable of transporting up to 450 tonnes (990,000 lb) of cargo per trip to Mars following propellant-refill in Earth orbit.[30]”

    From: ‘ITS launch vehicle’ Wikipedia
    At: https://en.wikipedia.org/wiki/ITS_launch_vehicle#Interplanetary_spaceship

    Yikes! Both the modified core of the SLS (the “Lunar-Mars Lander monstrosity –thing”) and the ITS (another “Lunar-Mars Lander monstrosity –thing”) might land about the same mass on Mars!

    What a coincidence!

    Life is sometimes pretty funny!

    • James

      And:

      The hydrolox RS-25 has sea a level Isp of 366 seconds.
      The hydrolox RS-25 has a vacuum Isp of 452 seconds.
      The hydrolox RS-25, if vacuum optimized, might have an Isp of over 470 seconds.

      The methalox Raptor has an expected sea level Isp of 334 seconds.
      The methalox Raptor has an expected vacuum Isp of 361
      The methalox Raptor, if vacuum optimized, has an expected Isp of about 382 seconds.

      Isp is an important rocket engine performance indicator.

      “Specific impulse (usually abbreviated Isp) is a measure of the efficiency of rocket and jet engines. By definition, it is the total impulse (or change in momentum) delivered per unit of propellant consumed[1] and is dimensionally equivalent to the generated thrust divided by the propellant flow rate.”

      And, “A propulsion system with a higher specific impulse uses the mass of the propellant more efficiently in creating forward thrust, and in the case of a rocket, less propellant needed for a given delta-v, per the Tsiolkovsky rocket equation.[1][3] In rockets, this means the engine is more efficient at gaining altitude, distance, and velocity.”

      From: ‘Specific impulse’ Wikipedia
      At: https://en.wikipedia.org/wiki/Specific_impulse

      “LOX and liquid hydrogen, used in the Space Shuttle orbiter, the Centaur upper stage of the Atlas V, Saturn V upper stages, the newer Delta IV rocket, the H-IIA rocket, and most stages of the European Ariane 5 rocket.”

      And, “However, liquid hydrogen does give clear advantages when the overall mass needs to be minimised; for example the Saturn V vehicle used it on the upper stages; this reduced weight meant that the dense-fueled first stage could be made significantly smaller, saving quite a lot of money.”

      From: ‘Rocket propellant’ Wikipedia
      At: https://en.wikipedia.org/wiki/Rocket_propellant#Advantages

      I know some long-haul truck drivers and they get excited when talking about the fuel efficiency of various large trucks.

      Perhaps rocket engineers planning on doing lots of efficient long-haul cargo missions to the Moon and across Cislunar Space and across our Solar System also get excited about propellant Isp.

      Cheers!

      • Oh for cryin’ out loud, specific impulse is not the only design criteria for rocket engines.
        In the real world where engineers work, trade-offs and compromises must be made in order to arrive at a practical design.

        Liquid hydrogen has a horrible energy density, requiring relatively huge tankage to hold the low-density fluid, and it must be kept at only ≈20 deg above absolute zero to remain liquid. This is why liquid methane makes a good compromise between hydrogen or storable propellants. And of course, methane is easy to synthesize from the Martian CO2 atmosphere.

        Of course “practical” and “compromise” mean nothing to someone who lives in a child’s world where Thunderbirds” is a documentary.

    • you accept that the ITS will someday be reusable over vast distances and in diverse environments?

      I’ve written nothing about ITS on this forum, there you go making things up.
      But ok, the SpaceX ITS is a two stage system, and yes, the upper stage will be reusable in diverse environments, while the booster stage will be reusable in the stage one launch vehicle environment. I personally think the ITS upper stage should be broken up into at least two specialized vehicles, but I digress.

      evolving RS-25 rocket engines have an excellent chance of being reusable over vast distance and diverse environments
      Again…NO. NO chance. What part of “hydraulic head pressure” of over 800 tons of liquid propellant don’t you understand? That’s a rhetorical question, you don’t seem to understand anything.

      The RS-25 can not be restarted in space, it’s a specialized launch vehicle engine not suitable in any way-shape-or-form for upper stage use.

      Look, you could take the Coyote 5 liter V8 engine out of my Fort truck, enlarge it, take out the pistons, crankshaft, cams, heads, and everything, then put in a 10 stage axial compressor, jet fuel burner cans, exhaust nozzle, a 130-inch bypass fan -and call it the “Ford Coyote 5 liter V8 Wide Body Jet Airliner Engine” -but that would be ridiculous.
      About as ridiculous as expecting the US government (or anyone) to spend 10s of billions of dollars to totally remake the RS-25 SSME into an upper stage engine for your equally ridiculous SLS core stage spaceship thing.

      If need be, buy a modified New Glen heat shield…
      New Glen is a launch vehicle, a rocket booster, it does not NEED a “heat shield”, like the Falcon-9 stage one, the New Glen Booster doesn’t go fast enough to require a heat shield for reentry. You do not seem to understand the concepts of velocity and kinetic energy, or why launch vehicles have stages, and why those stages are design optimized for different flight regimes.

      You don’t grasp the very basics of this stuff, but I’m sure that won’t stop you from posting more endless copy ‘n past nonsense or your sick hate screeds on Spudis’s blog.

      • James

        Blue Origin’s New Glenn is desgned to have a second stage that should eventually be reusable and have a heat shield.

        And Blue Origin seems to be planning on building a reusable launcher even much bigger than the New Glenn because they want to help NASA and international folks build a permanent ISRU colony on the Moon.

        Yep, “methane is easy to synthesize from the Martian CO2 atmosphere” and water, but maybe it is not so useful for Lunar missions. We should get some rovers going around taking samples of the volatiles in Lunar polar craters to get a better understanding of exactly what resources are available and where they are located.

        Which place are we going to commercially mine first, the Moon or Mars?

        Does Mars or the Moon hold the most accesable resources that would be quite useful in reducing human spaceflight risks and costs in developing Cislunar Space and doing lots of beyond Cislunar Space missions?

        Or are humans going to follow some super expanded, costly, and environmentally damaging space program that requires everything to be hauled up from the Earth?

        You claim, “This is why liquid methane makes a good compromise between hydrogen or storable propellants”.

        However, the Blue Origin folks who built the nifty hydrolox BE-3 seem to see things differently.

        “The engine is being used on the Blue Origin New Shepard suborbital rocket, for test flights which began in 2015.[2] The engine is under consideration by United Launch Alliance (ULA) for use in a new second stage, the Advanced Cryogenic Evolved Stage, in ULA’s Vulcan orbital launch vehicle with first flight in the 2020s.”

        And, “In January 2013, the company announced the development of the Blue Engine-3 BE-3, a new liquid hydrogen/liquid oxygen (LH2/LOX) cryogenic engine. The engine was originally announced to produce 440 kN (100,000 lbf) thrust, with initial thrust chamber tests planned for mid-February 2013 at NASA Stennis.[5]”

        And, “By December 2013, Blue Origin updated engine specifications following engine tests conducted on test stands at ground level, near sea level. This demonstrated that the engine could produce 490 kilonewtons (110,000 lbf) of thrust at full power, and could successfully throttle down to as low as 110 kilonewtons (25,000 lbf) for use in controlled vertical landings if needed for that purpose on particular launch vehicles.[7] The final engine specifications, released in April 2015 following the full test phase, included a minimum thrust of 89 kilonewtons (20,000 lbf), an even wider throttling capability by 20 percent than the preliminary numbers, while maintaining the previously released full power thrust spec.[8]”

        And, “For Blue Origin’s future orbital launch vehicle they are developing a variant of the BE-3 for use in upper stages. As of November 2015, the engine was projected to have a vacuum thrust of 670 kilonewtons (150,000 lbf).[15] An extendable nozzle for BE-3U is under development in early 2016.”

        From: ‘BE-3’ Wikipedia
        At: https://en.wikipedia.org/wiki/BE-3#cite_note-nsj20130117-5

        Clearly Blue Origin did not buy your “‘practical’ and ‘compromise'” methalox engineering for their new hydrolox BE-3 engine that they intend “for use in upper stages”.

        I do like that part about an “extendable nozzle for BE-3U”.

        Let’s also put an “extendable nozzle” on the RS-25 and try to get an Isp of well over 470 seconds.

        And let’s do an ITS type of engineering compromise on the core of the evolved SLS.

        “As of September 2016, SpaceX has identified two spacecraft that will also play the upper stage role on each Earth-away launch: Interplanetary spaceship and the ITS tanker. Both spacecraft are the same physical external dimensions: 49.5 m (162 ft)-long and 12 m (39 ft)-diameter (17 m (56 ft) across at the widest point. Both are powered by six vacuum-optimized Raptor engines, each producing 3.5 MN (790,000 lbf) thrust, and will have three lower-expansion-ratio Raptor engines to be used for in-space maneuvering as well as during descent and landing to allow for reuse on future launches.”

        From: ‘ITS launch vehicle’ Wikipedia
        At: https://en.wikipedia.org/wiki/ITS_launch_vehicle

        Wow! The ITS upper stage needs nine raptor engines…

        Ok!

        The future evolved and modified graphene core of the SLS could have nine highly efficient hydrolox rocket engines!

        Let’s have three ground started evolved extendable nozzle RS-25s and six ground started extendable nozzle BE-3Us that are used when launching from Earth, the Moon, or Ceres.

        To land on the Moon, Ceres, or Earth we’ll start up the six extendable nozzle BE-3Us first. And, then six seconds later, we can start the three evolved extendable nozzle RS-25s.

        Those three RS-25s remind me of something…

        It is a pity about Mars. Our new Super Duper Environmentalist President sought in 2029 and tonight finally got a hundred year UN ban on human and robotic orbital and surface missions to Mars due to the risk of environmental contamination and damage to any potential Martian bacteria.

        Of course, American AI androids are hinting that the President is not a real Earthling and that Mars has some secret and well-hidden colonies of super intelligent micro-nauts that are from far beyond our Solar System. Our AI android folks are pretty smart and so lots of other Americans are beginning to wonder if they are right…

        Cheers!

        • quoting myself:
          You do not seem to understand the concepts of velocity and kinetic energy, or why launch vehicles have stages, and why those stages are design optimized for different flight regimes

          edit:
          You do not seem to understand the concepts of velocity and kinetic energy, or why launch vehicles have stages, and why those stages are design optimized for different flight regimes

          Spend less time skimming Wikipedia, and pick up a book for a change; the internet is a poison pill for guys like you.

  • James

    Perhaps those trying to sell methalox and fantasies of super costly Mars colonies “do not seem to understand the concepts of velocity and kinetic energy”.

    “Spend less time” drinking snake oil mixed with pseudo Martian pixie dust, “and pick up a book” on hydrolox “for a change;” that pseudo Martian pixie dust and snake oil mixture “is a poison pill for guys like you.”

    Cheers!

  • James

    For those interested in the real world of SLS RS-25 hydrolox powered missions:

    “For example, ULA (a Boeing-Lockheed alliance) held a workshop in February with NASA and industry officials to talk about developing the “cislunar marketplace” (cislunar referring to the space between Earth and the moon). Perhaps private space companies could learn how to harvest ice on the moon, manufacture rocket fuel with it, and set up refueling stations in orbit.”

    And, “‘Rocket fuel is very heavy,’ says Burns. ‘You don’t want to take it off the Earth because it costs 10 times as much to take it off the Earth as it does to lift off the moon. So if you have the technology to develop this on the moon then store it in cislunar space with cryogenic tanks, then vehicles that are coming and going in the inner part of the solar system—going to asteroids, going to the moon, and going to Mars—will get their fuel from these cryogenic fuel depots.'”

    From: “Here Is the Trump Transition Team’s Big Plan for NASA
    Popular Mechanics met with astrophysicist Jack Burns, a member of President Trump’s NASA transition team, to learn a little about what we can expect from the agency moving forward.”
    At: http://www.popularmechanics.com/space/moon-mars/a26025/nasa-future-jack-burns-moon-mars-sls-orion/

    • James

      Maybe the sellers of methalox fantasy flights to super costly Martian Colonies should give up on their endless silly anti-SLS/International Orion and anti-Moon rants and their equally goofy love of super risky, costly, and slow and unhealthy methalox powered flights in Galactic Cosmic Radiation rich deep space flights.

      They could move into the RS-25 SLS real world of hydrolox and head out faster across all of our Solar System!

      The Exploration Upper Stage (EUS) has a potential bigger relative, the 24 meter long EDS (Earth Departure Stage) for Block II of the evolved SLS.

      “SLS Block II EDS
      Length 24 metres (79 ft)
      Engines 3 J-2X
      Thrust 3,930 kilonewtons (880,000 lbf)
      Specific impulse 448 seconds (vacuum)
      Fuel LH2/LOX”

      From: ‘Earth Departure Stage’ Wikipedia
      At: https://en.wikipedia.org/wiki/Earth_Departure_Stage

      Maybe swap in 6 or more Blue Origin extendable nozzle hydrolox BE-3Us for the “3 J-2X” engines on that potential SLS EDS.

      The SLS/International Orion space transportation system is funded, is being built, has some highly efficient reusable hydrolox engines, has broad national and international support, and is evolvable in lots of useful and interesting directions including:

      “An additional beyond-LEO engine for interplanetary travel from Earth orbit to Mars orbit, and back, is being studied as of 2013 at Marshall Space Flight Center with a focus on nuclear thermal rocket (NTR) engines.[64] In historical ground testing, NTRs proved to be at least twice as efficient as the most advanced chemical engines, allowing quicker transfer time and increased cargo capacity. The shorter flight duration, estimated at 3–4 months with NTR engines,[65] compared to 6–9 months using chemical engines,[66] would reduce crew exposure to potentially harmful and difficult to shield cosmic rays.[67][68][69][70] NTR engines, such as the Pewee of Project Rover, were selected in the Mars Design Reference Architecture (DRA).”

      From: ‘Space Launch System’ Wikipedia
      At: https://en.wikipedia.org/wiki/Space_Launch_System

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