Shuttle fuel burn

[ShotOne]

I've been quoted a statistic that the Space Shuttle uses 96.2% of its fuel to get enough the first foot off the ground. How true is this?

[criticalmass]

It takes about eight minutes for the liquid oxygen and liquid hydrogen in the main tank to be used up. The solid rocket boosters have a shorter burn time.

Simple logic suggests the figure you quote is somewhat in error.

[PAXfips]

Rather 9.6%?
Let alone the SSRBs (sidemounted boosters) ran for some over 90sec providing the majority of thrust and are jettisoned at ~45km over ground.
The main engine operates from the external tank, carrying 600t of LOX and 100t LH2. LOX was pumped at 1200kg/s, LH2 at 200kg/s.

Doesnt count up to 96% for the "first foot" for me.

[DaveReidUK]

Simple logic suggests the figure you quote is somewhat in error. Besides, if it was true, they would just build the launch site a foot higher. :O

[Dufo]

It took around 8 seconds from ingnition of three shuttle engines until liftoff. This time allowed systems check and after start checklist (just kidding) to be done until ingniton of solid boosters. Including first foot of ascent it took 9 seconds at most, which out of 500 total seconds of powered main engines flight represents less than 2% of fuel. Thrust/fuel flow was, excluding few seconds of reduction from 104% to 72% during initial ascent due to max dynamic pressure and last few seconds during which max g load was limited to 3g, constant.

[ACMS]

I love it when you talk dirty Dufo :p

[GordonR_Cape]

The Saturn V (Apollo moon rocket launches) were somewhat different from the Space Shuttle, with very gradual initial acceleration (seen in contemporary footage). Most of the thrust is used to overcome the sheer weight of the rocket, and only once the propellant begins to burn off, does this lead to increasing acceleration (given a constant thrust). This process might be part of the 'myth' that you quote.

The Wikipedia entry for Saturn V rocket data (https://en.wikipedia.org/wiki/Saturn V#Soviet_N1.2FL3) indicates initial acceleration (my calculation) of 2m/s^2 vs 5m/s^2 for the Shuttle (https://en.wikipedia.org/wiki/Space_Shuttle#Specifications).

There is a nice video showing altitude and velocity vs time https://www.youtube.com/watch?v=F0Yd-GxJ_QM for the entire first stage Saturn V burn (less than 3 minutes total), which consumes more than 4/5 of the rocket's launch weight. Just before stage cutoff, the acceleration is around 20 times that at liftoff. The second stage then goes though another period of gradual acceleration as its propellant burns off.

[PAXfips]

In all gory detail from STS-135:

00...07...30...11:19:16 AM...Orbiter access arm retraction
00...05...00...11:21:46 AM...Launch window opens
00...05...00...11:21:46 AM...Hydraulic power system (APU) start
00...04...55...11:21:51 AM...Terminate LO2 replenish
00...04...00...11:22:46 AM...Purge sequence 4 hydraulic test
00...04...00...11:22:46 AM...IMUs to inertial
00...03...55...11:22:51 AM...Aerosurface profile
00...03...30...11:23:16 AM...Main engine steering test
00...02...55...11:23:51 AM...LO2 tank pressurization
00...02...35...11:24:11 AM...Fuel cells to internal reactants
00...02...30...11:24:16 AM...Clear caution-and-warning memory
00...02...00...11:24:46 AM...Crew closes visors
00...01...57...11:24:49 AM...LH2 tank pressurization
00...00...50...11:25:56 AM...SRB joint heater deactivation
00...00...31...11:26:15 AM...Shuttle GPCs take control of countdown
00...00...21...11:26:25 AM...SRB steering test
00...00...07...11:26:39 AM...Main engine start (T-6.6 seconds)
00...00...00...11:26:46 AM...SRB ignition (LAUNCH)

(last value speed in mph)
11:26:46 AM...T+00:00...LAUNCH
11:26:57 AM...T+00:11...START ROLL MANEUVER.........................927
11:27:04 AM...T+00:18...END ROLL MANEUVER.........................1,002
11:27:18 AM...T+00:32...START THROTTLE DOWN (72%).................1,200
11:27:35 AM...T+00:49...START THROTTLE UP (104.5%)................1,432
11:27:47 AM...T+01:01...MAX Q (744 psf)...........................1,650
11:28:49 AM...T+02:03...SRB STAGING...............................3,627
11:28:59 AM...T+02:13...START OMS ASSIST (1:23 duration)..........3,743

full ascent (and other data along):
Spaceflight Now | STS-135 Shuttle Report | Ascent Timeline (http://spaceflightnow.com/shuttle/sts135/fdf/135ascent.html)

[sharmatt]

(nt)

Besides, if it was true, they would just build the launch site a foot higher. :O

:D:)

[Toryu]

(last value speed in mph)
11:26:46 AM...T+00:00...LAUNCH
11:26:57 AM...T+00:11...START ROLL MANEUVER.........................927
11:27:04 AM...T+00:18...END ROLL MANEUVER.........................1,002
11:27:18 AM...T+00:32...START THROTTLE DOWN (72%).................1,200
11:27:35 AM...T+00:49...START THROTTLE UP (104.5%)................1,432
11:27:47 AM...T+01:01...MAX Q (744 psf)...........................1,650
11:28:49 AM...T+02:03...SRB STAGING...............................3,627
11:28:59 AM...T+02:13...START OMS ASSIST (1:23 duration)..........3,743

927mph in 11s - talk about going places! :}

Those figures have to be taken in context:
That mph-value is an inertial speed. It's the speed of the Orbiter's (STS') center of mass rotating around Earth's center. Hence the inertial velocity of 914mph at standstill at T-00s.
Was the launchpad at either pole, the inertial velocity would be 0mph.
It's at maximum at the Equator - for obvious reasons.

The actual (aerodynamic/ relative to Earth's surface) velocities are given in the upper table as "velocity (e)".

That gives us 126mph after T+10s, which is a whole lot more reasonable.

[Spanner Turner]

Bare in mind the MPH speed figures quoted at the lower part of PAXfips post are "inertial" speeds. I.E cumulative speed of the vehicle + earth rotational velocity from west to east.
Before launch the vehicle is travelling 914.4 MPH due to the earths rotation.

If you open the link posted by PAXfips and look at the top section, the 6th column from the left will give you the orbiters "local" speed and the 7th column the "inertial" speed. e.g at 10 secs after liftoff "local" speed is 126.8 MPH and after 30 secs speed is 495.7 MPH.

:ok:

[Spanner Turner]

Toryu too fast.

An amazing machine through and through and no doubt much missed! :{

[FLX/MCT]

As far as I know once the SSRBs were ignited they could no longer be stopped for obvious reasons.
Were there any contingency scenarios in place for the case that only one booster started successfully?

[Two's in]

Were there any contingency scenarios in place for the case that only one booster started successfully?

There was a "Circle to Land" procedure if only one booster fired.

[Amadis of Gaul]

There was a "Circle to Land" procedure if only one booster fired.

I'd love to hear more about this. I understood there was essentially no contingency at all during ascent.

[Rwy in Sight]

There were a number or runways around the world as alternate if the lift-off does not go according the the plan. However the window of time to be used is very small.

[Dont Hang Up]

A brief foray into Wikipedia reveals the following

https://en.wikipedia.org/wiki/Space_Shuttle_abort_modes

It would seem that the return to land is only possible after SRBs are jettisoned.

Regarding the original question, I would be prepared to believe that 96.2% of the fuel load is required to carry the fuel load. It is just conceivable that that sort of statistic could be misrepresented by careless journalism as "96.2% just to get the the vehicle off the ground"!

[SincoTC]

There was a "Circle to Land" procedure if only one booster fired.

Hmm, if one SRB failed to ignite then I'm pretty certain that could only be a cartwheel to certain total loss of crew and vehicle :eek:

[pattern_is_full]

Regarding the original question, I would be prepared to believe that 96.2% of the fuel load is required to carry the fuel load.Exactly my thought. Thanks for pre-empting me!

The original quote is an idiomatic metaphorical analogy - to "get off the ground" (i.e. function to achieve the design goal of putting x amount of payload into Earth orbit), the shuttle needed fuel to move the payload, and fuel to move that fuel, and fuel to move the fuel that moves the fuel....to the point that 96.2% of the fuel is used just to move other fuel, over the entire climb and acceleration.

"Tankering" in the extreme.

[MG23]

Were there any contingency scenarios in place for the case that only one booster started successfully?

Yes. It was called 'crash and burn'.

If I remember correctly, the SRBs had three independent igniters, and only needed one to set them alight. So it was one of the less likely failure scenarios, but disastrous if it did happen. The SRBs were bolted to the pad with explosive bolts, but the bolts weren't strong enough to overcome the SRB thrust if the other didn't ignite (that would just introduce another catastrophic failure mode).

[Peter47]

Sorry this post is about the Saturn V which have studied in more detail than the Space Shuttle.

According to the Apollo 11 press kit the Saturn / Apollo combination burnt 37.2% of its propellent in the first 81 seconds at which stage it has reached 13,220m & 802m/s. (Bear in mind that it started at 409 m/s owing to the earth's rotatation.)

Hr.Min.Sec..Event......................................Altit ude....Velocity.... .Range...Time...Velocity....Altitude....Range...Distance.... ..Mass
............................................................ ...............ft............ft/s............nm.........s.........m/s............m.........km..............km.............kg
.....0..00.0..First motion.......................................183....1,341... ........0.0.......0.0........409.............56..........0.0 ...........0.1.......2,938
.....1..21.0..Maximum dynamic pressure........43,365.....2,637.........2.7......81.0...... ..804......13,218..........5.0.........14.1.......1,856
.....2..15.0..Centre engine cutoff..................145,600.....6,505........24.9....135 .0.....1,983......44,379........46.1.........64.0.......1,15 1
.....2..40.8..Outboard engine cutoff..............217,655.....9,031........49.6....160.8.. ...2,753......66,341........91.9.......113.3.......1,084
.....2..41.6..First stage separation................219,984.....9,065........50.2....1 61.6.....2,763......67,051........93.0.......114.6.......... 658
.....2..43.2..Second stage ignition................221,881.....9,059........51.3....163 .2.....2,761......67,629........95.0.......116.6
.....3..11.5..Aft interstage jettison.................301,266.....9,679........87.0....19 1.5.....2,950......91,826......161.1.......185.5
.....3..17.2..LET jettison...............................315,001.....9,778.... .....94.3....197.2.....2,980.....96,012.......174.6......199 .3
.....7..39.8..Centre engine cutoff..................588,152...18,762......600.0.....459. 8....5,719...179,269....1,111.2....1,125.6
.....9..11.4..Outboard engine cutoff..............609,759...22,747......885.0.....551.4... .6,933...185,855....1,639.0....1,649.5
.....9..12.3..Second stage separation...........609,982...22,757......888.0.....552.3.. ..6,936...185,923....1,644.6....1,655.0..........175
.....9..15.4..Third stage ignition....................610,014...22,757......888.4..... 555.4....6,936...185,932....1,645.4....1,655.8
...11..40.1..Third stage first cutoff................617,957...25,562...1,425.2.....700.1.. ..7,791...188,353....2,639.5....2,646.2
...11..50.1..Parking orbit insertion................617,735...25,568...1,463.9.....710. 1....7,793...188,286....2,711.1....2,717.7
...44..14.8..Third stage reignition.................650,558...25,554...3,481.9....... ..0.0....7,789...198,290.....6,448.5...6,451.5
2:50..03.1..Third stage second cutoff.......1,058,809...35,563...2,633.6.....348.3..10,840. ..322,725.....4,877.4...4,888.1
2:50..13.1..Translunar injection................1,103,215...35,539...2,605.0.....35 8.3..10,832...336,260.....4,824.5...4,836.2

https://www.hq.nasa.gov/alsj/a11/A11_PressKit.pdf

Does anyone know why the first stage used kerosene rather than liquid hydrogen & oxygen in the way that shuttle does given that the latter has a higher specific impulse so less fuel would have been required? Presumably the rocket would just have been too large or were there other reasons?

[tdracer]

Does anyone know why the first stage used kerosene rather than liquid hydrogen & oxygen in the way that shuttle does given that the latter has a higher specific impulse so less fuel would have been required? Presumably the rocket would just have been too large or were there other reasons?
The Saturn V was really pushing the state of the art - and the F1 first stage engines were several times more powerful than anything previously attempted (1.5+ million pounds of thrust). While H2 engines have a huge specific impulse advantage, kerosene is much easier to deal with since it's not cryogenic and H2 is far less dense that kerosene of the same impulse, so the first stage would have gotten even bigger. Also, the first stage is, as noted, only there for the first 80 seconds. It basically was cheaper, easier, and safer to simply make the first stage heavier and use kerosene, rather than 'optimize' it by using H2. The upper stages, being far more weight critical, used H2 (I don't recall the specific number, but one additional pound on the third stage added something like 20 lbs. of propellant needed to the first stage).
It's basically the same reason they used the Solid Rocket Boosters on the shuttle (solid boosters have worse specific impulse than kerosene).


BTW, my understanding was not only was the failure of an SRB to light catastrophic, they needed to light within a few milliseconds of each other to avoid catastrophic loads on the shuttle stack. Given how hard it is to light ammonium perchlorate (AP) based propellant, I always thought that a bit scary (we use a very similar AP based propellant in hobby rockets, and 'misfires' (failure to light) are not uncommon). :rolleyes:

[joema]

"Were there any contingency scenarios in place for the case that only one booster started successfully?"

No. The maximum survivable L/R SRB ignition differential was about 100 milliseconds -- beyond that and the structure would fail.

The SRBs were originally planned for "thrust termination ports" in the nose which could be blown off and equalize the thrust allowing SRB jettison before solid fuel depletion. Later studies showed it would require 20,000 lbs of additional structural reinforcement, or about 1/3 total payload capacity, so this was dropped.

However the SRBs are best viewed as large pyrotechnic devices. The shuttle and Apollo/Saturn before it had many "criticality 1" pyros that simply had to work. E.g, the Lunar Module ascent and descent stages had no release couplings, they were cut apart with pyros, including a rocket-powered guillotine which cut through the plumbing. There was no backup; it simply had to work.

The often-stated safety measure of shutting down a liquid-fueled engine is less comprehensive than first appears. If the Saturn V had a single engine failure during the first 14 seconds of flight, thrust:weight ratio dropped below 1:1 and it would fall back onto the pad. All five engines had to work perfectly, at least that long.

The shuttle SRBs had sufficient steering authority and thrust to enable a liftoff and abort if all three SSMEs failed, however this was only possible (theoretically) after a structural upgrade in the mid-to-late 1990s.

The shuttle abort options were dramatically improved following STS-51L (Challenger) in 1986. Previously there were long "black zones" during the ascent where no abort was possible due to various structural, aerodynamic or guidance/control factors.

The contingency abort improvments can be seen in figure 7 and 8 in the 2011 AIAA document "Space Shuttle Abort Evolution"(now reproduced in the Wikipedia article Space Shuttle Abort Modes): http://tinyurl.com/jm8zelv

The SRBs themselves were vastly improved from a reliability standpoint after Challenger. The changes were comprehensively discussed in a book by Allan J. McDonald.

Another risk of shutting down a liquid engine is the possibility of sensor-induced spurious shutdown. This happened on STS-51F in 1985 and came within seconds of a second SSME shutdown which would not have been survivable, since that was before they had bailout. A quick-thinking flight controller had them override the redline limits, which placed the SSMEs in "open loop" mode without any safety monitoring. Fortunately they made it to a low orbit; ironically that was also Challenger: https://www.youtube.com/watch?v=0Rz82mf01Yo

[Acrosport II]

DaveReidUK
Besides, if it was true, they would just build the launch site a foot higher.

Nice one,..

Yeah, By that logic, If they had more Irish on the Space Shuttle program, they could have seriously saved some money on their fuel bill!

Employ More Irish!

[PAXfips]

Thanks for the additions on speed, I wasnt really aware of that.

About the return options, just check SFN for the missions and all what joema said.

Just plain power: https://www.youtube.com/watch?v=JBU78AxAm2c

I miss the Shuttle.

[Check Airman]

Interesting series of podcasts on the website Omega Tau (http://omegataupodcast.net/):

http://media.libsyn.com/media/omegataupodcast/omegatau-43-flyingTheSpaceShuttle.mp3

http://traffic.libsyn.com/omegataupodcast/omegatau-132-theSpaceShuttle-I.mp3

http://traffic.libsyn.com/omegataupodcast/omegatau-133-theSpaceShuttle-II.mp3

He does a great job of getting into some technical details with guests who have first hand knowledge. In one part, he discusses contingency plans for problems during launch.

His Concorde (http://traffic.libsyn.com/omegataupodcast/omegatau-166-flyingTheConcorde.mp3) interview is highly informational as well.

[VinRouge]

Interesting also the SS had a self destruct option controlled by the range safety officer.

Not a nice job to flick the switch.

Space Shuttle Destruct Switch - NASA Prepared to Blow Up Discovery (http://www.popularmechanics.com/space/a3232/4262479/)

[Toryu]

The SRBs were actually blown up once: During Challenger's final flight, when the SRBs flew unguidedly along after the ET explosion/ Orbiter breakup.

Challenger didn't really explode: It was accelerated away from the collapsing fuel-tank at a rate higher than it's structure could sustain. It broke up under aero-loads.

[joema]

Here is a shuttle abort simulation done around STS-26. The audio is the actual abort; the events are sync'd to video of a successful launch. This simulated abort included a cascade of failures which resulted in inability to RTLS, with ocean ditching the only option:

https://www.youtube.com/watch?v=HGyiLv6mPbg

The simulated failures include:

- AC bus phase 3 failure
- Right engine failure
- MDM (multiplexer/demultiplexer) failure, this is the link between the main computers and secondary computers
- IMU (Inertial Measurement Unit) accelerometer failure. IMUs are the primary on-board navigation method
- Left engine failure
- Center engine failure
- loss of voice communications

Callout terminology used:

"Abort RTLS" -- Return to Launch Site abort, where vehicle pitches around and heads back to Kennedy
"MDM FF2 down at the MDM" -- multiplexer/demultiplexer failure which cuts off communications between computers and other subsystems.
"I/O reset picking up" - Emergency keyboard command, attempt to restore flight avionics after a critical failure
"IMU #1 accelerometer problem" -- Shuttle navigates by three redundant IMUs (Inertial Measurement Units). One failed due to bad accelerometer.
"Deselect/reselect IMU #1" -- try to get IMU back on line by rebooting it.
"Dump in progress" -- emergency firing of the orbital maneuvering engines to lighten the vehicle.
"We're in 601" -- computers are running program 601, designed for RTLS abort
"two engine out blue" -- a procedural call indicating they've lost two engines early in the ascent
"need the pushbuttons" -- automatic engine shutdown upon propellant depletion is unreliable, must manually shut down engines
"any chance of getting this one back?" -- Is shuttle too far downrange for a return to launch, or must it ditch in the ocean
"have you got a predicted IP?" -- Can the ocean landing impact point be calculated?
"602" -- flight computers are running program 602 (RTLS abort subroutine for reentry). RTLS subroutine is used even for a bailout over the ocean.
"need to reduce alpha" -- during reentry the nose is pitched too high, if uncorrected would over-G the vehicle
"He's in the pullout" -- Orbiter is in atmospheric reentry, must pull up, but stay within thermal/G limits.
"Select the lower left antenna" -- Orbiter has several flush-mounted antennas on the upper fuselage which communicate with TDRSS satellites. Request they manually select one to restore comm.

[Dont Hang Up]

The often-stated safety measure of shutting down a liquid-fueled engine is less comprehensive than first appears. If the Saturn V had a single engine failure during the first 14 seconds of flight, thrust:weight ratio dropped below 1:1 and it would fall back onto the pad. All five engines had to work perfectly, at least that long.

However I think Saturn V had much more realistic abort options over the Shuttle. A solid fuel rocket motor was attached directly to the nose cone that contained the command module. The command module could therefore be dragged clear by this (to safe parachute and splashdown height).

Once the emergency escape rocket was jettisoned the vehicle was high enough for the whole Command/Service module to power itself clear of the main structure.

The lunar lander was of course another story. Scarily lacking in any contingency a long long way from home.

[joema]

"I think Saturn V had much more realistic abort options over the Shuttle. A solid fuel rocket motor was attached directly to the nose cone that contained the command module. The command module could therefore be dragged clear by this (to safe parachute and splashdown height).

Once the emergency escape rocket was jettisoned the vehicle was high enough for the whole Command/Service module to power itself clear of the main structure."

This is basically true, but the presumed safety factor of launch abort can be misleading. The launch escape tower is jettisoned early in the 2nd stage, which leaves the only a "mode II" abort option of separating CSM, which had a meager 0.32-to-1 thrust-to-weight ratio. The separation would have been very slow.

This almost happened on Apollo 13, when a severe longitudinal oscillation happened on the center engine of the 2nd stage. This was briefly shown in the movie. In the actual mission, later analysis showed the vibration was so severe it nearly caused structural failure. The center engine was vibrating with a fore/aft stroke of 3 inches -- and it's mounted on a large metal cruciform beam, similar to a skyscraper I-beam.

13 Things That Saved Apollo 13, Part 5: Unexplained Shutdown of the Saturn V Center Engine - Universe Today (http://www.universetoday.com/62672/13-things-that-saved-apollo-13-part-5-unexplained-shutdown-of-the-saturn-v-center-engine/)

Saturn V Also Suffered Engine Launch Anomalies : Discovery News (http://news.discovery.com/space/history-of-space/saturns-engine-failures-121011.htm)

Instrumentation revealed the engine vibration peaked at 68 g. The cutoff was essentially a lucky event that probably saved the vehicle and possibly the crew. Had the 2nd stage broken up it's unclear if the sluggish CSM could have accelerated away safely.

While Apollo 13 was the only manned mission of that series that experienced a significant booster anomaly, the unmanned Apollo 6 also had severe pogo problems that would have probably triggered an abort. In that case it happened on stage 1, so they theoretically could have used launch escape, although it was at a very high dynamic pressure.

On the Skylab Saturn V launch, pyros failed to separate the interstage skirt between stage 1 and 2. On a manned mission this was officially an abort scenario since the additional weight would have prevented reaching orbit. Officially this would have been a tower abort, although I'm not sure a 12g abort would have been needed in that more gradual situation.

A non-obvious criteria governing Apollo abort rules is the reentry heating and g load must be survivable. The CM had a L/D ratio of about 0.3:1, so while it had some lift to modulate g and heat loads, there were otherwise possible abort trajectories that exceeded g or thermal limits. The goal was avoid these but I vaguely recollect there were conditions where abort initiation was possible but it was not survivable.

The Apollo Launch Escape System weighed about 8,000 lbs, vs the CM roughly 13,000 lbs -- 62% of vehicle weight. At a significant payload cost, it did provide an abort option over the first 25% of the ascent until it was jettisoned.

[MG23]

The lunar lander was of course another story. Scarily lacking in any contingency a long long way from home.

It was basically designed to minimize the risk of failure, rather than to have backups if it did fail. The ascent stage engine, for example, was about as simple as you can make a rocket engine; if I remember correctly, it was pressure-fed and hypergolic, so just a couple of tanks, a couple of valves, and a rocket nozzle.

NASA actually did propose to carry an emergency backup for the LEM ascent stage on later missions which would stay on the Moon for longer periods with a greater risk of failure. Basically just a rocket engine, a couple of fuel tanks that would be filled with enough fuel from the LEM to reach low orbit, and space for two astronauts to stand in their space suits.

[joema]

"It was basically designed to minimize the risk of failure, rather than to have backups if it did fail. The ascent stage engine, for example, was about as simple as you can make a rocket engine; if I remember correctly, it was pressure-fed and hypergolic, so just a couple of tanks, a couple of valves, and a rocket nozzle."

That is correct, and I would not describe the LM as lacking contingencies. The design philosophy was extreme reliability through simplicity -- even at a significant cost of performance.

The LM used batteries instead of fuel cells to improve reliability, the engines had no turbopumps, no ignitors, no engine gimbal in the ascent stage, no regenerative cooling -- it was as simple as could be made. The propellant valves were double redundant and the plumbing had multiple contingency paths.

The LM was the first digital fly-by-wire vehicle, and like modern aircraft, flight control was critical. The LM had dual-redundant fly-by-wire guidance & navigation computers hand-built and tested with extreme rigor, and the backup computer had independent software developed by a separate company to avoid a generic failure.

If both LM flight control computers failed, there was a manual reversion mode including a non-computerized analog pathway from the hand controller to separate redundant solenoid propellent actuators on each RCS thruster. Activating this required only a full deflection "hard over" on the hand controller to trigger the alternate path.

There was a practiced contingency procedure to ascend from the lunar surface and achieve orbit using no computers and no inertial platform whatsoever. It involved using charts, a stopwatch and making stepped pitch changes to align window etchings with the lunar horizon. Astronaut Gene Cernan said he achieved this in the simulator and felt it was possible.

[Amadis of Gaul]

Interesting also the SS had a self destruct option controlled by the range safety officer.



I wonder what research they did to conclude that a Space Shuttle heading for "populated areas" in a million flaming pieces would do less damage than if it were still in more or less one piece.

[Max Angle]

I wonder what research they did to conclude that a Space Shuttle heading for "populated areas" in a million flaming pieces would do less damage than if it were still in more or less one piece. It would have been destroyed well before it started pointing anywhere other than out to sea, as someone pointed out above the Range Safety Officer has an unenviable job in the event of a launch going badly off course.

[Jwscud]

Mike Mullane's excellent autobiography touched on this. RSOs did not socialise with astronauts to avoid clouding their judgement.

I thoroughly recommend "Riding Rockets" to anyone interested in the shuttle or aviation in general.

[AC560]

Mike Mullane's excellent autobiography touched on this. RSOs did not socialise with astronauts to avoid clouding their judgement.

I thoroughly recommend "Riding Rockets" to anyone interested in the shuttle or aviation in general.

Interesting to note that originally there was a caution light in the shuttle showing when the RSO armed the self destruct. This was to indicate they should eject, when they took the seats out of the Columbia after the first four flights they left the light in (and in subsequent versions). Not a warning light you would like to see go off I imagine. :sad::sad::sad:

[bratschewurst]

The Saturn V (Apollo moon rocket launches) were somewhat different from the Space Shuttle, with very gradual initial acceleration (seen in contemporary footage). Most of the thrust is used to overcome the sheer weight of the rocket, and only once the propellant begins to burn off, does this lead to increasing acceleration (given a constant thrust). This process might be part of the 'myth' that you quote.

The first stage of the Saturn V generated around 7.6 million lbs of thrust at lift-off, with a total stack weight of around 6.8 million lbs (figures varied slightly from Apollo 9 to 17). The STS SRBs generated 5.6 million lbs thrust and the three SSME's generated another 1.2 million lbs, pushing a total stack mass of around 4.5 million lbs. So it's not surprising that the STS was accelerating much more smartly at lift-off; its thrust-to-weight ratio was quite a bit better at launch.

On the other hand, toward the end of the Saturn V first-stage burn, the center engine would be cut in order to limit acceleration to under 4Gs. By this point the vehicle had burnt off close to 4.5 million lbs; ie it was less that half the weight it had been at lift-off 2.5 minutes previously.

[bratschewurst]

It was basically designed to minimize the risk of failure, rather than to have backups if it did fail. The ascent stage engine, for example, was about as simple as you can make a rocket engine; if I remember correctly, it was pressure-fed and hypergolic, so just a couple of tanks, a couple of valves, and a rocket nozzle.

It was also the only stage of the Saturn stack that had a non-gimbaling engine, making it even simpler; all directional control was done by the reaction control system, which was almost as simple and, of course, had been tested on the descent.

On the other hand, IIRC it was also the only engine in the Saturn stack that couldn't be tested before launch, as the fuels were so corrosive that the engine could only be used once before needing extensive rebuilding. So lighting it was always a bit of a nail-biter; unlike every other engine in the vehicle, if it didn't work the first time, the astronauts died.

[bratschewurst]

This almost happened on Apollo 13, when a severe longitudinal oscillation happened on the center engine of the 2nd stage. This was briefly shown in the movie. In the actual mission, later analysis showed the vibration was so severe it nearly caused structural failure.

The second stage thrust structure was apparently rated to 15Gs, which was somewhere between 1/2 and 1/4 of what it experienced during Apollo 13. It's entirely possible that this was actually the closest the Apollo program came to losing a crew after the Apollo 1 fire.

Prior to this flight, POGO was viewed as more of a problem for astronaut comfort and their ability to function than as a potential structural failure issue, which is why the Saturn engineers were willing to live with it. After Apollo 13, they put it a POGO suppressor and the problem did not re-occur. The Shuttle was designed with POGO suppression from the get-go.

[bratschewurst]

"Does anyone know why the first stage used kerosene rather than liquid hydrogen & oxygen in the way that shuttle does given that the latter has a higher specific impulse so less fuel would have been required? Presumably the rocket would just have been too large or were there other reasons?"

The Saturn V was really pushing the state of the art - and the F1 first stage engines were several times more powerful than anything previously attempted (1.5+ million pounds of thrust). While H2 engines have a huge specific impulse advantage, kerosene is much easier to deal with since it's not cryogenic and H2 is far less dense that kerosene of the same impulse, so the first stage would have gotten even bigger. The J-2 engines (using cryogenic H2/O2) were a huge advance over anything that had been done before using H2 as a fuel; even so, they only generated 110,000 lbs of thrust vs 1.5 million for the F-1. The SSMEs were, in turn, a huge advance over the J-2s, and even they only produced 400K lbs or so. If Apollo had needed H2/O2 engines for the first stage, it likely never would have made it to the moon.

If I understand matters correctly, it's also true that the specific thrust advantage of H2/O2 engines mattered less at lower speeds and denser atmospheres.

[riff_raff]

The shuttle SRB's provided most of the thrust at launch. But the RS-25 liquid engines were used because they could be throttled. The RS-25 liquid engines were started about 3 seconds prior to igniting the solid boosters. This allowed the liquid engine turbopumps to spool up and produce thrust. If you watch a video of the RS-25 engines starting up, you'll see an initial spray of liquid propellants just before the engines ignite.

[Uplinker]

And the engines needed to be throttled down soon after launch to avoid exceeding max Q or max structural dynamic pressure loads on the STS.

To do this, they took the shuttle main engines down to 67% thrust in the lower atmosphere and then back up to 104% when the atmospheric density reduced sufficiently.

The solid boosters were more sophisticated than I imagined. They each had an APU and vectored thrust nozzle.

Have a look at this superb footage of a launch. The Shuttle main engines start first - to make sure they are all healthy - before the boosters are started. You can see (and hear) the whole STS angle over with the force, and then the boosters are lit.

Awesome !!

https://m.youtube.com/watch?v=Lq_shHu4lAs

[Bushfiva]

If you watch a video of the RS-25 engines starting up, you'll see an initial spray of liquid propellants just before the engines ignite.

If you're referring to around T-7, that's the sound suppression system, which is a water blanket designed to protect the shuttle stack from damage due to noise reflecting from the concrete.

[MG23]

To do this, they took the shuttle main engines down to 67% thrust in the lower atmosphere and then back up to 104% when the atmospheric density reduced sufficiently.

The SRBs also 'throttled down' at that point by a change in the shape of the propellant at that stage of the burn.

Have a look at this superb footage of a launch. The Shuttle main engines start first - to make sure they are all healthy - before the boosters are started. You can see (and hear) the whole STS angle over with the force, and then the boosters are lit.

Pedantically, the boosters were lit when the shuttle returned to the vertical after the SSME ignition pushed the stack forward and it swung back.

[Max Angle]

Also check out this film, also available as an iPad app (NASA Ascent), 45 minutes of gorgeous slow motion launch photography from the numerous cameras on and around the pad narrated by two guys who ran the operation.

https://www.youtube.com/watch?v=vFwqZ4qAUkE

One great geeky detail of the SSME start process was that the engine bell vibrated quite markedly as the engine lights up. The flight position of the two lower engines (no. 2 and 3) placed the bells very close together and there was a chance they could collide during the start so they were held apart during light up and then gimballed back into position once they are up and running. Easily visible on close ups of engines during launch.

Another vote for Riding Rockets by Mike Mullane, great behind scenes look at the operation, warts and all. One thing that is very clear from the book is that the astronauts themselves were under no illusions about how dangerous and marginal the vehicle was and the risks they were taking with every launch.

[No Fly Zone]

I think GordonR missed his boat or flight. Space Cadet?
The original question said NOTHING about the Saturn V. Not a single word. Focus, please...:p

[YRP]

Besides, if it was true, they would just build the launch site a foot higher.

:)

Invoice for keyboard on its way.

[Tcraft41]

Somewhere I found that the last check before the SRB's are lite off all three of the main engines test their gimbaling and only after being successful on all three engines do they fire the SRB's or do an abort if any one fails.

[Uplinker]

The whole American space program was mind blowing. I have visited KSC twice, and I spent ages just looking at the Saturn V and all the engines in awe.

Another awesome fact is that the Shuttle's rate of climb is 2000 feet per second. I will write that again, 2000 feet per second.

The Airbus I fly can climb at 2000 - 2300 per minute with normal loading........

[joema]

"The Space Shuttle definitely accelerates more quickly from the launch pad (twice as quickly as the Saturn V)"

The more rapid liftoff for the shuttle was due to higher initial thrust-to-weight ratio, but this changed vs the Saturn V during the ascent.

From a thrust standpoint, the Saturn V sea-level liftoff thrust on Apollo 15 was 7.823 million lbf (34.8 MN), which increased with altitude to a peak of 9.18 million lbf (40.8 MN) at T+135 seconds. The engines were not throttled; this was due to increasing nozzle efficiency as ambient pressure dropped:

Image:SaturnVThrust2.jpg - from the Schools Wikipedia (http://www.cs.mcgill.ca/~rwest/link-suggestion/wpcd_2008-09_augmented/images/745/74550.jpg.htm)

The shuttle liftoff thrust was about 6.78 million lbf (30.16 MN). The SRBs underwent significant throttling over their approx. 120 sec firing time, as can be seen in this graph. This throttling was not dynamic but designed in by controlling propellant surface area:https://en.wikipedia.org/wiki/Space_Shuttle_Solid_Rocket_Booster#/media/File:Srbthrust2.svg

Like the Saturn V, the shuttle SRB and SSME engines improved with altitude, but they both throttled back around Max Q, roughly T+60 sec. Despite improving nozzle efficiency with altitude, the SRB throttle schedule probably meant the peak vehicle thrust was at liftoff not at altitude like the Saturn V. So comparing peak thrust to peak thrust, it was about 6.78 million lbf vs about 9.18 million lbf.

Despite the slower start, at higher altitudes the Saturn V accelerated more rapidly, reaching a peak of 3.8 g just before 1st stage cutoff.

Of course the ultimate goal is delivering a payload. On Apollo 15 the Saturn V delivered a payload of 140,930 kg (310,697 lbs) to low earth orbit. The heaviest shuttle payload was about 23,586 kg (52,000 lbs).

It would have taken six shuttle launches to deliver to LEO the same useful payload as a single Saturn V.

[tdracer]

The shuttle liftoff thrust was about 6.78 million lbf (30.16 MN). The SRBs underwent significant throttling over their approx. 120 sec firing time, as can be seen in this graph. This throttling was not dynamic but designed in by controlling propellant surface area:https://en.wikipedia.org/wiki/Space_...Srbthrust2.svg (https://en.wikipedia.org/wiki/Space_Shuttle_Solid_Rocket_Booster#/media/File:Srbthrust2.svg)Getting the desired time/thrust curve is quite the science for solid rockets. Unlike the simple 'end burning' black powder rockets, most AP based solid rocket motors are 'core burning' and are ignited at the front. The shape of that core cavity determines the time/thrust relationship. IIRC, the Shuttle boosters had a 'star' shaped core to get the desired characteristics.

It would have taken six shuttle launches to deliver to LEO the same useful payload as a single Saturn V. The shuttle was a tremendous technological achievement, but it failed miserably at it's primary goal of reducing the cost of inserting payload into orbit by at least an order of magnitude. By most measures, even accounting for inflation, the cost of placing a pound of payload into orbit was not meaningfully different between the Space Shuttle and Saturn V. About the only meaningful capability we gained with the shuttle was the ability to bring large objects back from orbit.

I find it both amusing and pathetic that NASA is now spending billions of dollars to recreate the capability we already had 45 years ago with the Saturn V :ugh:

[Uplinker]

Wow.

So does that mean that the Shuttle and its fuel tank could have used the Saturn V first stage instead of the boosters?

I presume the boosters were much easier and vastly cheaper to manufacture?

I didn't know the booster thrust varied during the launch - very clever.

[Amadis of Gaul]

Thanks to all who have posted above. There's a huge expertise here I didn't know existed.



Indeed. The amount of expertise around here is truly overwhelming at times.

[joema]

"Does that mean that the Shuttle and its fuel tank could have used the Saturn V first stage instead of the boosters?

I presume the boosters were much easier and vastly cheaper to manufacture?"

The Saturn V S-IC first stage was considered as a shuttle booster:

https://joema.smugmug.com/Aerospace/Saturn-V-Shuttle-Booster/i-TmrdjpS/A

However the initial goal was a fully-reusable two-stage vehicle. Unfortunately the optimum staging velocity for lowest total mass is around Mach 12, which means a reusable first stage would essentially be a hypersonic transport in the 3-4 million pound gross weight class. This would have entailed extremely high development cost and risk. Former shuttle program manager Robert F. Thompson said even had they been given the money, he didn't think it would have been possible.

This iteratively drove the design process to an expendable booster (the above Saturn S-IC was one of many concepts), which then led to an orbiter with external propellant, which then led to an expendable tank with semi-reusable solid rocket boosters.

See "The Space Shuttle Decision": The Space Shuttle Decision: Chapter 6 (http://www.nss.org/resources/library/shuttledecision/chapter06.htm#intro)

It was understood early on the shuttle would likely not be economical to operate. This was covered in the 1972 GAO report to Congress: http://archive.gao.gov/f0302/096542.pdf

There was also a RAND study in that period which had similar statements about projected per-flight cost, also reviewed in the Jenkins book, "Space Shuttle: History of the National Space Transportation System", p. 173. It was known in the early 1970s even at 60 (!) flights per year, the shuttle payload cost per pound would only be about 38% cheaper than expendable boosters. If that flight rate could not be achieved it would be (and was) more expensive.

The vehicle was never designed to reach that flight rate, e.g, the maximum production rate of external tanks using three shifts at the Michoud facility was only 24 tanks per year.

The shuttle was an amazing vehicle but in hindsight there was no possible design in the early 1970s that would have permitted achieving the conflicting performance, safety, cost and flight rate goals.

[Amadis of Gaul]

I get the sense everybody knew this was a bad idea, but did it anyway. Gotta love the government.

[tdracer]

I get the sense everybody knew this was a bad idea, but did it anyway. Gotta love the government.
Not so much a bad idea as ahead of its time - the technology just wasn't ready yet. I was only a teenager when NASA abandoned the fully reusable 'two stage' shuttle concept - but even then I knew that basically meant throwing out the baby with the bathwater.

NASA has long been accused of being more interested in "sexy" than in functionality, and the shuttle was "sexy". There have long been advocates of "Big Dumb Rockets" - simple, (relatively) cheap, not man rated so reliability was not as critical. One proposal was for a Saturn 1B sized rocket based around a single F-1 engine, million plus pound launch weight that could put the same payload as the shuttle into orbit but for much lower cost (there was even talk of recovery and reuse of the F-1 engines but I don't know how viable it was). But that NASA wasn't interested because such a rocket was boring and the shuttle was sexy :rolleyes: Although to be fair, boring doesn't get as much funding as sexy :E

My personal vision has long been for a two stage shuttle - horizontal takeoff, the first stage being air breathing using hydrocarbon liquid fuel and some combination of turbojets/ramjets/scramjets, and a pure rocket H2/O2 orbiter. Of course it would cost a large fortune to develop, but per launch costs would basically be fuel and maintenance. Sadly it won't happen in my lifetime. Neither the money or the technology are there to make something like that viable.

It still seriously ticks me off that all that engineering that went into the Saturn V was lost - I still consider the Saturn V/Apollo (along with the associated moon landings) to be the greatest technological achievement of the 20th century and we literally threw that away.

[Luther Sebastian]

One way of thinking of the acceleration the crew of the orbiter experience is to imagine doing 0-60 in just under two seconds. But continuously for 8.5 minutes...

[bratschewurst]

The fundamental problem with the space shuttle program was the same basic problem that NASA had post-Apollo 11; there simply was no appetite on the part of either the public nor the political leadership for any potential follow-on missions that would justify the kind of funding required for successful completion of those programs.

As the document referenced in post #57 made clear, the design of the shuttle system - a re-usable vehicle with a pretty small payload, launched with the aid of two massive solid-rocket boosters and a disposable tank - was chosen because it was the cheapest to develop at a time when NASA budgets were being pared back significantly in real terms. They knew at the time it wouldn't be the cheapest to operate, but they weren't given the money to develop a better or more sustainable program.

To give one example of the choices forced on NASA after Apollo 11, the last three moon missions were cancelled, even though the hardware was built, and even though the cost of launching those missions was a tiny fraction of what had been spent on the space program, which had been driven entirely by the goal of getting to the moon. Yet NASA cancelled almost 1/3 of the moon missions - although hundreds of billions in current dollars had been spent to develop the capability to launch those missions - simply to free up some cash for shuttle development, because they weren't allocated the money to do both.

The ultimate irony of the shuttle program was that it was most designed to build, service and re-supply a space station - which was, due to budgetary constraints, abandoned pretty early on in shuttle development. It became a machine with the sole mission of flying people into space to do - what?

No doubt NASA made some bad decisions post-Apollo. But the fundamental decision that crippled the space program was the decision by the political leadership and the American people not to have an Apollo-scale space program at all.

[MG23]

No doubt NASA made some bad decisions post-Apollo. But the fundamental decision that crippled the space program was the decision by the political leadership and the American people not to have an Apollo-scale space program at all.

History was rather more complex than usually made out. See, for example http://history.nasa.gov/SP-4221/ch9.htm

[riff_raff]

If you're referring to around T-7, that's the sound suppression system, which is a water blanket designed to protect the shuttle stack from damage due to noise reflecting from the concrete.No, I wasn't referring to the blast of water below the engines used to dampen engine noise. Take a look at the engine nozzle on the left in this image. The light colored plume just right of the nozzle centerline is propellants dumping out of the combustion chamber that have not yet ignited.

http://www.spaceflightnowplus.com/hd/images/sts133film_ssmeHD.jpg

[bratschewurst]

History was rather more complex than usually made out. See, for example http://history.nasa.gov/SP-4221/ch9.htm
While I'm not clear precisely what point you are trying to make, I agree that the document you cite is well worth reading. As a case study on decision-making about pushing the limits of engineering in a complex political environment, with serious financial constraints, it is unsurpassed by anything else I've read.

[bratschewurst]

Have a look at this superb footage of a launch. The Shuttle main engines start first - to make sure they are all healthy - before the boosters are started. You can see (and hear) the whole STS angle over with the force, and then the boosters are lit.
https://m.youtube.com/watch?v=Lq_shHu4lAsAnother revealing look at a shuttle launch is here: https://www.youtube.com/watch?v=C0OmJFFQp50 (https://www.youtube.com/watch?v=C0OmJFFQp50)

If you look at 5:07, you will see just how violent the ignition of the solid rocket boosters was.

Also extremely interesting is this video, which is part 2 of about 30 minutes of 400 fps video of the launch from various engineering cameras: https://www.youtube.com/watch?v=M0L5CBrqE2U (https://www.youtube.com/watch?v=M0L5CBrqE2U)

Go to 9:20 to see just how rough was the thrust that the SRBs produced; essentially the thrust pulsed about around 7 hz. SRBs were not an elegant technology.

[bratschewurst]

Originally posted by JOEMA:
Quote:
Originally Posted by MG23
"It was basically designed to minimize the risk of failure, rather than to have backups if it did fail. The ascent stage engine, for example, was about as simple as you can make a rocket engine; if I remember correctly, it was pressure-fed and hypergolic, so just a couple of tanks, a couple of valves, and a rocket nozzle."

That is correct, and I would not describe the LM as lacking contingencies. The design philosophy was extreme reliability through simplicity -- even at a significant cost of performance.One of the most interesting things I learned from reading about Apollo 13 was that NASA's design philosophy assumed that structures didn't - or wouldn't - fail. So structures, unlike systems, generally did not have backups or much redundancy. The reason was simple - the weight penalty would have been prohibitive. So they designed the structures and non-redundant systems to be a simple as possible.

There were multiple navigation systems on both the CSM and the LM, and multiple sources of electrical power on both as well. But there was only one SPS engine (albeit with two sets of valves for the hypergolic fuels) and only one descent engine and ascent engine on the LM.

They also tried to design the systems so that they could use what redundancy was inherent - being able to use the LM engine to power the entire stack, for example, as was done in Apollo 13, or more broadly being able to use the LM as a lifeboat for those portions of the mission when it was attached. But that fell well short of full redundancy. If the O2 tank on Apollo 13 had exploded when the LM was on the moon, or on return to earth, it would have caused the loss of the crew.

[Amadis of Gaul]

My personal vision has long been for a two stage shuttle - horizontal takeoff, the first stage being air breathing using hydrocarbon liquid fuel and some combination of turbojets/ramjets/scramjets, and a pure rocket H2/O2 orbiter. Of course it would cost a large fortune to develop, but per launch costs would basically be fuel and maintenance. Sadly it won't happen in my lifetime. Neither the money or the technology are there to make something like that viable.


Perhaps that's because such a setup is not viable period. To me, the whole idea of trying to combine more than one type of vehicle in one (e.g. a "spaceplane", a tilt-rotor, a "flying car", a half-track etc) is that you end up with something that combines most of the DISadvantages of both with only a few of the advantages. In this example the "orbiter" still has wings, some kind of landing gear, some sort of an aerodynamic control system for re-entry, none of which it needs at any time OTHER than re-entry. It also has to carry all the "spaceship stuff" which it does not need during the re-entry. In other words, at any given time in the mission, good half the mass of the vehicle is dead weight that's sitting there doing nothing, but still has to be carried. Is this not the very definition of inefficiency? It's hard enough to build a good airplane or a good spaceship, but to build something that's both is, in my opinion, well-nigh impossible.

All this is before we even get to your first stage vehicle which will ostensibly be this fairly large, cumbersome, expensive airplane that will only be good for one thing.

[tdracer]

Perhaps that's because such a setup is not viable period. To me, the whole idea of trying to combine more than one type of vehicle in one (e.g. a "spaceplane", a tilt-rotor, a "flying car", a half-track etc) is that you end up with something that combines most of the DISadvantages of both with only a few of the advantages. In this example the "orbiter" still has wings, some kind of landing gear, some sort of an aerodynamic control system for re-entry, none of which it needs at any time OTHER than re-entry. It also has to carry all the "spaceship stuff" which it does not need during the re-entry. In other words, at any given time in the mission, good half the mass of the vehicle is dead weight that's sitting there doing nothing, but still has to be carried. Is this not the very definition of inefficiency? It's hard enough to build a good airplane or a good spaceship, but to build something that's both is, in my opinion, well-nigh impossible.


Not sure what you're getting at - all vehicles contain some level of compromise to maximize their usefulness. A commercial jetliner carries around landing gear, flaps, brakes, etc. which make up a significant portion of the aircraft mass but are worthless for 95% of the flight. But the aircraft itself is worthless if it can't takeoff or land.
A spacecraft that can't re-enter and slow for landing is similarly of very limited value.
The Space X landable booster has to carry considerable extra weight for the landing struts and fuel in order to soft land - but the alternative is for the entire booster to be disposable. Isn't a single use spacecraft the real definition of inefficiency?

[Jonno_aus]

Space Shuttle Systems 101 - More Than You Ever Needed To Know About The Space Shuttle Main Engines (http://www.interspacenews.com/FeatureArticle/tabid/130/Default.aspx?id=2130)


Some interesting stuff. Couple things caught my untrained and knowledge-less eye :

The power settings during flight are controlled automatically but if necessary the pilot can manually throttle the engines with a thrust lever located to the left of his seat, this is the same lever used to operate the speed brake during landing. The commander also has a similar lever for the speed brake but it is not capable of throttling the engines.


Wonder what fancy program was written to stop the throttles from also operating the speed brake if the pilot had to manually take over after launch? Obviously not a problem later on on approach if no rockets attached. And limiting having too many levers etc on the flight deck. Interesting nontheless.


At this point the vehicle will begin a roll, pitch and yaw maneuver that places it on its back (crew heads down) as it tracks the appropriate flight path for the desired orbit


From what I've read that lasts just over 5 mins in that position until it rolls back with the crew then 'heads up'. Must be a disconcerting feeling with the g-forces, vibrations, noise and I guess some blood draining to the head for some of that time.


Couldn't pay me enough to be an astronaut. I guess they get paid well..

[Amadis of Gaul]

Not sure what you're getting at - all vehicles contain some level of compromise to maximize their usefulness.

Indeed. It's exactly the level of compromise that's the issue. Landing gear, flaps and brakes comprise only a very small percentage of an airline's weight, unlike what we're talking about here.

[MG23]

Wonder what fancy program was written to stop the throttles from also operating the speed brake if the pilot had to manually take over after launch?

There were separate programs (or program modes) for launch and landing, so, with the throttle lever in auto mode, I guess it would depend on which program you were running.

Looks like there was also a manual override to switch between throttle and speed brake:

HSF - The Shuttle (http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/avionics/gnc/sb.html)

From what I've read that lasts just over 5 mins in that position until it rolls back with the crew then 'heads up'. Must be a disconcerting feeling with the g-forces, vibrations, noise and I guess some blood draining to the head for some of that time.

Gravity will affect the shuttle and crew at the same rate, so it shouldn't matter. Most of the effects the crew felt would come from thrust and drag.

And most of the vibration went away when the SRBs separated.

Couldn't pay me enough to be an astronaut. I guess they get paid well..

I think the astronauts who gave a talk I went to in the 90s said it was in the region of $40-50,000 a year at that time. So decently middle-class, but not exactly City banker levels. I don't think anyone did it for the money!

[Peter47]

I think that US Apollo astronauts were on service pay - the typical rank was Colonel or Naval Captain and they probably got less than half what a long haul airline captain got. Then again they could retire after twenty years.

You could earn rather more from being an astronaut. I understand that at some stage during the Apollo (or It may have been the Gemini) programme astronauts shared the proceeds of a very lucrative contract with, I think, Time magazine. Many went on to front advertising campaigns (Armstrong Chrysler, Shepard beer, etc.)

I also saw that Buzz Aldrin was charging a five figure sum for a public speaking engagement around ten years ago. The trouble is that whilst the man in the street might be able to name the Apollo 11 astronauts, how many could name those who flew the later missions. I don't know how much Tim Peake could get as a public speaker but he won't be making a fortune as an Army major.

[Jonno_aus]

Fascinating stuff. Also found this from Canadian astronaut Commander Chris Hadfield :

Launch is immensely powerful, and you can truly feel yourself in the centre of it, like riding an enormous wave, or being pushed and lifted by a huge hand, or shaken in the jaws of a gigantic dog. The vehicle shakes and vibrates, and you are pinned hard down into your seat by the acceleration. As one set of engines finishes and the next starts, you are thrown forward and then shoved back. The weight of over 4 Gs for many minutes is oppressive, like an enormous fat person lying on you, until suddenly, after 9 minutes, the engine shut off and you are instantly weightless. Magic. Like a gorilla was squishing you and then threw you off a cliff. Quite a ride :)

He also noted that it’s not possible to pass out during the launch, because you are being pushed into space while lying on your back, so your blood doesn’t end up draining out of your brain. Also, he said it takes about 15 seconds to go from a sunny day to complete darkness.


Brave indeed. And I'll happily admit now I'm not that brave and I'd only do it for the money. :E

[chksix]

John Young who has ridden almost all vehicles to space: "-Smooth as glass" :E

[joema]

"My personal vision has long been for a two stage shuttle - horizontal takeoff, the first stage being air breathing using hydrocarbon liquid fuel and some combination of turbojets/ramjets/scramjets, and a pure rocket H2/O2 orbiter. Of course it would cost a large fortune to develop, but per launch costs would basically be fuel and maintenance."

Winged HTOL air breathing shuttles (esp SSTO) have long been the ultimate dream. Superficially it's an alluring concept: just make it like an airliner. The U.S. spent billions of $ on the National Aerospace Plane (NASP) with this goal, and the UK is still backing Skylon.

Unfortunately the ugly reality is orbital velocity requires gigantically higher kinetic energy -- KE = 1/2*m*v^2, and every element in the design hinges on this. An air breathing first stage doesn't provide much performance advantage, yet entails huge financial cost and development risk.

When you calculate what kind of mothership is needed to launch an orbiter having a meaningful useful payload, it's something like a ramjet-powered XB-70 with 4x the gross weight. Then on top of that you need a self-powered orbiter capable of accelerating from Mach 5 to Mach 25.

NASP required multiple breakthroughs on many levels -- unobtanium materials, active cooling of the structure, scramjet propulsion which no wind tunnel can test, etc. A detailed account of this development is in the on-line publication "The Hypersonic Revolution": http://tinyurl.com/h8advzk

Getting into orbit via airbreathing propulsion has been called "getting to space the hard way". Unlike the casino game of craps, you don't get extra payoff for getting there "the hard way". Rather you want to get there the easiest, simplest way possible. Given current technology, that's probably some kind of a rocket.

[riff_raff]

Here's a very interesting document from NASA describing the Transoceanic Abort Landing (TAL) sites for the Shuttle. (https://www.nasa.gov/centers/kennedy/pdf/167472main_TALsites-06.pdf)It was a big effort preparing all of the emergency landing sites for each Shuttle launch.

A TAL could be declared at T+2:30. But neither a Return To Launch Site (RTLS) or TAL could begin until the SRB's were jettisoned. This chart shows the abort conditions for loss of one or more SSMEs (https://upload.wikimedia.org/wikipedia/commons/5/5a/ShuttleAbortPre51L.png).

A RTLS abort procedure was interesting since it could require a hair-raising maneuver called a Powered Pitch Around (PPA). A PPA involved pitching the Orbiter 180deg at an altitude of >400kft so that the SSME thrust could be used to propel it back to the launch site for landing.

Here is an image of the Shuttle's abort mode selector switch:

https://upload.wikimedia.org/wikipedia/commons/3/32/Space_Shuttle_abort_panel.jpg

[joema]

"A TAL could be declared at T+2:30. But neither a Return To Launch Site (RTLS) or TAL could begin until the SRB's were jettisoned. This chart shows the abort conditions for loss of one or more SSMEs."

That chart only applies before the 1986 Challenger accident. Afterward there were many enhancements which enabled intact aborts or at least bailout aborts. For higher-inclination launches (e.g, ISS missions) these also included East Coast Abort Landings (ECAL) at various sites along the U.S. east coast.

The abort improvments can be seen by comparing figure 7 and 8 in the 2011 AIAA document "Space Shuttle Abort Evolution": http://tinyurl.com/jm8zelv

[chksix]

It was said somewhere that had a TAL happened it would have set an unbeatable record for quickest crossing of the Atlantic.

[Jonno_aus]

https://m.youtube.com/watch?v=8gBBGb_isAs


I see they're given a couple intructions along the way for 'abort boundaries' at 3:30. Guess they change pretty quickly too as it accelerates at ridicules speeds. :eek:

[tdracer]

Getting into orbit via airbreathing propulsion has been called "getting to space the hard way". Unlike the casino game of craps, you don't get extra payoff for getting there "the hard way". Rather you want to get there the easiest, simplest way possible. Given current technology, that's probably some kind of a rocket.

You'll note that I did specify that the technology isn't there yet. But as long as we continue to throw away a significant portion (or all) of a rocket every time we use it, space access will remain insanely expensive (currently between $2,600 and $10,000 per pound - and the shuttle was ~ $24,000/lb :eek:)

Something around 75% of the liftoff weight of an orbital capable rocket is oxidizer - get rid of that for the first stage and you can add a whole bunch of hardware to make the thing reusable while still reducing the overall vehicle size and weight. True, the current state of the art for Ramjet/Scramjets is no where near making this viable, but technology marches on. An air breathing launch vehicle that could carry an orbiter to ~120,000 ft./Mach 7 (or better) would mean the orbiter wouldn't need to be huge to carry the fuel to get into orbit.

As I noted, it's not going to happen in my lifetime - the money and technology are not there. Worse, the will to do things like going to the moon is long gone. But if we're ever going to get to the stage where traveling into space is as routine and commonplace as traveling cross country, that's what it's going to take.

[joema]

"as long as we continue to throw away a significant portion (or all) of a rocket every time we use it, space access will remain insanely expensive (currently between $2,600 and $10,000 per pound - and the shuttle was ~ $24,000/lb )

This is the standard argument; it has some validity but is misleading. A good example is the reusable shuttle cost about $10,000-$25,000 per lb to LEO. The expendable SpaceX Falcon Heavy currently costs under $1,000 per pound, and that was achieved with no reusability, no exotic airbreathing, no wings, etc.

In his presentation to the Augustine Commission, shuttle program manager John Shannon called reusability a myth. He explained when a vehicle is made in very small numbers, you cannot just shut down all manufacturing capacity after it's built. The vehicle and subsystems require ongoing R&D, fixes, testing, etc. Parts wear out, you have failures, design issues become apparent during use. You essentially have to keep the production line available, even if nothing is being manufactured, along with much of the associated industrial infrastructure. Then you must buy "one of" pieces to fix things which is expensive.

Airliners achieve lower costs through mass production, not just reuse. They also benefit from a smaller "standing army" of support personnel relative to the fleet size.

Any envisionable air breathing SSTO would be far more complex and expensive than the space shuttle. If built in small numbers, there's no mass production. It would be costly to operate, regardless of having wings and airbreathing propulsion.

"Something around 75% of the liftoff weight of an orbital capable rocket is oxidizer - get rid of that for the first stage and you can add a whole bunch of hardware to make the thing reusable while still reducing the overall vehicle size and weight."

An air breathing design would save oxidizer. However LOX costs about $0.01 (one penny) per pound. The LOX for a shuttle launch cost only about $14,000. You are right much of the liftoff mass of an orbital rocket is oxidizer. On the shuttle the tankage cost for that was maybe about 1/2 of the total $50 million ET, so eliminating the LOX would enable smaller, cheaper tankage.

OTOH the additional cost and mass of a hypersonic airbreathing design offsets the gain from losing the oxidizer. An air breather must fly a depressed ascent trajectory, like a shuttle reentry in reverse. You must have active cooling of the vehicle skin and wings, multiple propulsion systems, fuel, structure and thermal protection for those systems, etc. Since scramjets cannot reach orbit, conventional rocket propulsion must be used for the final ascent. All the mass, structure and thermal protection for that must be hauled to and from orbit.

...True, the current state of the art for Ramjet/Scramjets is no where near making this viable, but technology marches on.

This is true so there is always hope. The NASP research made significant progress in various areas. With today's more advanced CFD and material science it might be more achievable.

An air breathing launch vehicle that could carry an orbiter to ~120,000 ft./Mach 7 (or better) would mean the orbiter wouldn't need to be huge to carry the fuel to get into orbit."

This is not quite like it seems because of the v^2 term. Mach 7 is only 7.8% of orbital kinetic energy. Being at 120,000 ft actually helps relatively little since gravity and air drag losses to reach that altitude are not severe. If launched at that altitude and velocity the orbiter would still have to do over 80% of the work to reach orbit, yet a Mach 7 mother vehicle would be fantastically heavy and expensive. Optimum staging velocity for lowest total vehicle mass on a two-stage launcher is about Mach 10, and the total vehicle mass curve is very steep around that minimum. IOW deviate much from Mach 10 staging and the whole stack get a lot heavier, quickly.

A lot of smart people have worked on airbreathing launchers for many years. It is a compelling concept but in the real world it is much more difficult than first appears. The current trend is that optimization of conventional rocket technology through lean production and partial re-use are a more effective way forward.

[Amadis of Gaul]

It seems to me the words often attribute to Ed Henneman are very applicable here. Simplicate and add lightness.


Rockets seem to accomplish that.

[riff_raff]

NASA performed studies of a fully re-usable system (http://arc.aiaa.org/doi/abs/10.2514/6.1978-1469) just prior to the Shuttle program start. And much later on studied a Liquid Fly Back Booster (LFBB) (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19980231024.pdf) concept that would replace the SRBs.

As for reducing $/lb launch costs, it will be interesting to see how effective the StratoLaunch system (http://stratolaunchsystems.com/) works out.

[tdracer]

This is not quite like it seems because of the v^2 term. Mach 7 is only 7.8% of orbital kinetic energy.
Wrong, quite wrong. Rockets engines are effectively constant thrust devices, not constant power. Discounting aerodynamic drag, it takes exactly the same amount of thrust to accelerate a given mass from zero to 10,000 ft./sec. as it does to accelerate from 10,000 to 20,000 ft./sec.

It seems to me the words often attribute to Ed Henneman are very applicable here. Simplicate and add lightness.

Rockets seem to accomplish that.
1916 biplanes were quite simple and light (not to mention cheap). That doesn't mean they are preferable to a very expensive, complex 2016 airliner. Heck, you could make a commercial airliner much simpler and cheaper if you threw it away after every flight - but not exactly cost effective.

Lots of people thought Musk was nuts when he proposed to use a powered descent and landing to recover his first stage - but he's close to making it work and is on record as stating that, once perfected, it'll reduce orbital costs by at least an order of magnitude (and Space X is already one of the cheapest per pound). Yes, it adds cost, weight, and complexity, but he's a pretty sharp guy, with some pretty sharp people working for him and if can successfully undercut the competition by an order of magnitude on costs, all the others are going to be in a mad scramble to avoid being completely left out.

As long as we persist in throwing away most of our launch vehicles after a single use, space is going to remain prohibitively expensive. Of course, if we persist in leaving thousands of pieces of used rockets and other debris up there, low earth orbit may well become unusable in another 20 years or so...

[riff_raff]

While I agree that Mr. Musk is bright guy, and SpaceX has made progress in reducing the $/lb of commercial launch costs, your claim that SpaceX will lower the $/lb of commercial launches by "an order of magnitude" is not realistic. According to the SpaceX website, the $/lb price of an LEO Falcon 9 is over $2K, not including other costs such as insurance, etc.

Commercial space launch is a business, and launch services are priced based on what customers are willing to pay versus the competition.To propose that SpaceX would offer launches at 1/10 the price of their competitors ignores basic market economics.

[Amadis of Gaul]

A cheap and simple 1916 biplane may be more suitable for, say, cropdusting and barnstorming than a complex and expensive 2016 airliner, however.

In other words, to Caesar that which is Caesar's and to Bubba Ray that which is Bubba Ray's.

[joema]

"...your claim that SpaceX will lower the $/lb of commercial launches by "an order of magnitude" is not realistic. According to the SpaceX website, the $/lb price of an LEO Falcon 9 is over $2K, not including other costs such as insurance, etc...."

According to SpaceX's own web site, the launch price for a Falcon Heavy is $90 million, and it can put 116,845 lbs in LEO. That is $770 per pound:

Falcon Heavy | SpaceX (http://www.spacex.com/falcon-heavy)

This is important because any alternative launcher, whether existing or planned, expendable, reusable, air breathing, etc, must compete with that number. That is already cheaper by an order of magnitude than the space shuttle was. Compared to the Delta IV Heavy ($5,908 per lb to LEO), it is 7.76 times cheaper.

[tdracer]

While I agree that Mr. Musk is bright guy, and SpaceX has made progress in reducing the $/lb of commercial launch costs, your claim that SpaceX will lower the $/lb of commercial launches by "an order of magnitude" is not realistic
Just to be clear, that's not my claim, it's Mr. Musk's claim. There was an interview a few years ago where he stated his intent to get launch costs down to ~$10/lb. to LEO.

[joema]

"...An air breathing launch vehicle that could carry an orbiter to ~120,000 ft./Mach 7 (or better) would mean the orbiter wouldn't need to be huge to carry the fuel to get into orbit.""
"...because of the v^2 term. Mach 7 is only 7.8% of orbital kinetic energy..."
"Wrong, quite wrong. Rockets engines are effectively constant thrust devices, not constant power. Discounting aerodynamic drag, it takes exactly the same amount of thrust to accelerate a given mass from zero to 10,000 ft./sec. as it does to accelerate from 10,000 to 20,000 ft./sec."

As I stated, because KE=1/2*m*v^2, it takes tremendous kinetic energy to reach orbit. A Mach 7 reusable booster is easier to design but it makes the orbiter much heavier. This is because the remaining delta-V is a squared term, hence energy required falls disproportionately on the orbiter.

Since the development cost of similar aerospace vehicles tends to scale upward with gross mass, It is conceivable a TSTO design with a Mach 7 booster would cost more than one with a Mach 12 booster. This can be non-intuitive, but some relevant data and research are in this graph, taken from "Propulsion and Staging Considerations for an Orbital Sortie Vehicle (Stengel, 1987):

https://joema.smugmug.com/Aerospace/Staging-Velocity/n-rXFf6t/i-nKLmZSR/O

Original paper (PDF): http://www.princeton.edu/~stengel/Prop.pdf

A reusable air-breathing first stage does seem compelling and has been discussed and studied many times. The problem is it requires a very large, very heavy, very complex hypersonic airbreathing launch vehicle, whether that is Mach 7 or Mach 12. Maybe a lower, more achievable staging velocity is the answer, providing a lighter, less expensive launcher?

E.g, assume launch vehicle is twice the gross weight of an XB-70 (about 1 million lbs), and can reach Mach 5 @ 120,000 ft. Mach 5 is a good number because it is (barely) within ramjet limits so avoids the uncertainty, development risk and extra mass of scramjets.

Assume it can carry twice the XB-70 external payload, about 100,000 lbs. Given a Mach 5 running start, could a 100,000 lb winged manned orbiter reach orbit with any useful payload?

The X-15A-2 had a mass ratio of about 0.72. We'll assume technical progress improves this to 0.8. The XLR-99 engine had an ISP of about 256 sec. Assume ISP improves to about 314 sec. Rough guess at exhaust velocity: 3000 m/sec. Using the rocket equation, this gives:

delta V = Ve * ln (full mass/dry mass)
delta V = 3000 m/sec * ln (45000 kg / 9000 kg)
delta V = 4828 m/sec (10,800 mph)

This would give a final delta V (launch ship + drop ship) of roughly 6,528 meters/sec, far short of 8,000 m/sec orbital velocity.

Thus it appears a winged, manned orbiter could not reach orbit -- even with zero payload -- if boosted to Mach 5 by a ramjet-powered XB-70 successor with double the payload capacity.

Air-breathing winged launchers and winged orbiters look cool, and have long been romanticised. However in the real world they don't work that well. This was discussed by Henry Spencer in the article "Rockets, not air-breathing planes, will be tomorrow's spaceships":

https://www.newscientist.com/blogs/shortsharpscience/2009/03/rockets-not-air-breathing-plan.html

[tdracer]

As I stated, because KE=1/2*m*v^2, it takes tremendous kinetic energy to reach orbit. A Mach 7 reusable booster is easier to design but it makes the orbiter much heavier. This is because the remaining delta-V is a squared term, hence energy required falls disproportionately on the orbiter.No, you're missing basic physics. Acceleration is directly proportional to force, and in this case force and thrust are the same thing (i.e. acceleration is force/mass) Work (or power) is force times velocity (in other words, the faster a rocket goes, the more work it does even if the thrust says constant). Energy is proportional to V squared, but acceleration is still directly proportional to thrust.
I won't bother with the rest of your analysis because it's based on a flawed assumption.

[balsa model]

Commercial space launch is a business, and launch services are priced based on what customers are willing to pay versus the competition.To propose that SpaceX would offer launches at 1/10 the price of their competitors ignores basic market economics.
Right, you are. But let's stick to the launch costs; not the launch price tag. The cost is very much a component of engineering and hence more interesting in Tech Log.

About this claim with respect to return to launch pad
once perfected, it'll reduce orbital costs by at least an order of magnitude
So far, they "only" returned the 1st stage. It is arguably the easier one to return because its kinetic energy is lower. Also, the mass penalty of carrying recovery h/w is relatively lower. It's all about the mass fraction that you impact.
Given that the 2nd stage would have to start its recovery maneouver at around Mach 25, the weight of thermal protection alone will eat directly into the revenue payload. I don't have the maths for it but my gut feel is that it would net an *increase* in $/lb to orbit.
But... if (big if) the cost of the launching h/w is proportional to the engines, then: 1st stage=9 Merlins; 2nd stage=1 Merlin. A divide by 10, right here.
Perhaps the claim wasn't that far fetched, if we allow a few qualifiers.

One argument weighting against the "reusables" is that such approach cuts into production rate and thus raises the cost of the few launchers that are being build (reduced economies of scale, etc.). Being valuable means that the launch will now include loss insurance. Also, being more valuable, there will be an incentive to make them more wear and random failure proof, farther eating into our mass. All such insidious practicalities will erode the final $/lb.

Personally, I can't wait until they announce that they are flying a refurbished stage and tell us what it took.

[balsa model]

$10/lb
Sanity check:

1 lb in LEO => Mach 25 => 0.5 m v^2 = 12 MJ
barrel of crude oil = 6100 MJ

Hm.. Indeed, it must be all down to "shipping and handling" :)

[AC560]

One argument weighting against the "reusables" is that such approach cuts into production rate and thus raises the cost of the few launchers that are being build (reduced economies of scale, etc.).

3D printing though is really changing the economics for building many of these unique and complicated parts versus when the space shuttles were flying.

[joema]

"An air breathing launch vehicle that could carry an orbiter to ~120,000 ft./Mach 7 (or better) would mean the orbiter wouldn't need to be huge to carry the fuel to get into orbit."

...because KE=1/2*m*v^2, it takes tremendous kinetic energy to reach orbit. A Mach 7 reusable booster is easier to design but it makes the orbiter much heavier. This is because the remaining delta-V is a squared term, hence energy required falls disproportionately on the orbiter.

...Wrong, quite wrong. Rockets engines are effectively constant thrust devices, not constant power....No, you're missing basic physics. Acceleration is directly proportional to force, and in this case force and thrust are the same thing...I won't bother with the rest of your analysis because it's based on a flawed assumption.

Ironically a Saturn V stages at Mach 7. You can see the event in this photograph. The remaining stack is still gigantic and massive -- 1.48 million lbs. The Mach 7 assist did not translate into a "non huge" orbital vehicle: https://joema.smugmug.com/Aerospace/Saturn-V-at-1st-stage/n-csvsxt/i-dCpvGLS/O

Another example is the Space Shuttle, which at SRB sep is at Mach 4.5 and 150,000 ft. About 77% of ET propellent is remaining, so the remaining stack is about 2.2 million pounds. Mach 4.5 is a good number because it's achievable with ramjet propulsion at much lower cost and technical risk than scramjets.

A hypothetical air-breathing first stage which injected the shuttle orbiter plus ET to Mach 4.5 at 150,000 ft. would reduce the required ET propellant by 23%. The remaining orbiter stack would still be a gigantic vehicle weighing 2.2 million lbs. Why? Because of the remaining tremendous kinetic energy required to reach orbit.

So these examples illustrate the difficulty of an air-breathing first stage. One key reason is the highly non-linear affect of the kinetic energy V^2 term. Boosting to Mach 4.5 or Mach 7 doesn't help that much, yet would require a titanically expensive and complex winged air-breathing booster.

The non-linear effect of how kinetic energy affects this can be easily seen in the near-vertical curve on the left side of the staging velocity chart: https://joema.smugmug.com/Aerospace/Staging-Velocity/n-rXFf6t/i-nKLmZSR/O

This also explains why subsonic air launch has almost no benefit to orbiter mass. This was discussed in the AIAA paper "A Study of Air Launch Methods for RLVs
(Sarigul-Klijn, et al, 2001): http://tinyurl.com/z7zstrm

""Surprisingly, a typical straight and level subsonic horizontal air launch such as used by the X-15 research rocketplane does not result in any significant changes in the delta V requirement as compared to a baseline vertical surface launch."

There are groups still working on air-breathing SSTO and TSTO launchers such as Skylon and Bristol. I wish them well but I doubt they will be successful. Even if they were successful, they would still have to compete with SpaceX price/performance, while amortizing the huge development cost of airbreathing hypersonic vehicles:

Skylon: Reaction Engines Ltd - Space Access: SKYLON (http://www.reactionengines.co.uk/space_skylon.html)
Bristol SpaceBus: Spacebus | Bristol Spaceplanes (http://bristolspaceplanes.com/projects/spacebus/)

Air-breathing SSTO is not dumb -- the brilliant designer Tony duPont conceived the original NASP design and it first appeared possible. However DARPA recently re-evaluated this in light of technical growth and believes it is still not possible: http://www.nytimes.com/2014/10/21/science/25-years-ago-nasa-envisioned-its-own-orient-express.html?_r=0

NASP and follow-on research indicates scramjets and the associated penalties of multiple propulsion systems, active cooling, etc, cannot remotely reach orbit and are not good candidates for a 1st stage reusable launcher. It now appears the main application for hypersonic airbreathing propulsion is military -- long-range hypervelocity missiles.

However a reusable non-airbreathing 1st stage is possible for smaller payloads - the XS-1 is a good example: US Military Awards New Contracts for XS-1 Space Plane (http://www.space.com/30196-xs1-military-space-plane-boeing-contract.html)

The problem is even with the projected reuse savings of that booster, their payload cost per lb is already higher than SpaceX on the Falcon Heavy ($1,000 vs $770). It may have other advantages such as quick reaction and flexibility.

[riff_raff]

Right, you are. But let's stick to the launch costs; not the launch price tag. The cost is very much a component of engineering and hence more interesting in Tech Log.For commercial launches isn't the most important consideration the difference between the costs incurred by the business providing the launch and what the customer pays for the service? You cannot successfully operate a business that loses money on every transaction.

From what I have seen there is not enough difference in the technology or design of the current SpaceX Falcon 9 booster and other existing boosters to justify the claims of massive reductions in launch costs. The current expendable Falcon 9 uses fairly conventional materials and design. The only major difference seems to be that SpaceX employs a more efficient manufacturing process.

Can someone provide an explanation of why a launch on an expendable SpaceX Falcon 9 vehicle could be so much cheaper than other similar launchers?

[joema]

"From what I have seen there is not enough difference in the technology or design of the current SpaceX Falcon 9 booster and other existing boosters to justify the claims of massive reductions in launch costs. The current expendable Falcon 9 uses fairly conventional materials and design. The only major difference seems to be that SpaceX employs a more efficient manufacturing process....Can someone provide an explanation of why a launch on an expendable SpaceX Falcon 9 vehicle could be so much cheaper than other similar launchers?"

This is a good question, which many people have asked. There were initially lots of doubts that the cost reduction goals could be met. There are still doubts about whether the low prices are sustainable, and questions about how exactly SpaceX is achieving this. There are some general answers but SpaceX keeps many specific details confidential for competitive reasons.

Apparently the main answers are SpaceX launchers were designed from scratch with cost reduction as a primary goal, they do most things in house rather than subcontract, and their business organization is very lean. Here are some articles discussing it:

Air & Space: "Is SpaceX Changing the Rocket Equation?" History, Travel, Arts, Science, People, Places | Air & Space Magazine (http://www.airspacemag.com/space/is-spacex-changing-the-rocket-equation-132285884/?page=1)

Motherboard: "How Elon Musk Willed SpaceX Into Making the Cheapest Rockets Ever Created" How Elon Musk Willed SpaceX Into Making the Cheapest Rockets Ever Created | Motherboard (http://motherboard.vice.com/read/how-elon-musk-willed-spacex-into-making-the-cheapest-rockets-ever-created)

[MG23]

Can someone provide an explanation of why a launch on an expendable SpaceX Falcon 9 vehicle could be so much cheaper than other similar launchers?

Because it's designed to be cheap, rather than efficient. The space shuttle had the most efficient engines ever built at that time, which is one reason it was so expensive; because they were so complex, and operated beyond their original design thrust, they cost a lot to build, and they cost a lot to reuse. Another reason was because it required about 10,000 people to operate, whereas SpaceX uses around a tenth as many.

And reusability will massively reduce costs, if they can reuse a stage a few times with minimal refurbishment. If returning the first stage to the launch site cuts 1/3 off the payload, and you can reuse it only twice more without replacing or overhauling the engines, you've nearly halved the cost in $/lb to orbit.

[India Four Two]

Does anyone know why they didn't reorient the shuttle on the launch pad, instead of performing the roll manoeuvre?

[PAXfips]

Does anyone know why they didn't reorient the shuttle on the launch pad, instead of performing the roll manoeuvre?
53 Why does the shuttle roll just after liftoff? (http://stason.org/TULARC/science-engineering/space/53-Why-does-the-shuttle-roll-just-after-liftoff.html)

I would say that this explains it a bit with the bottom line: historical deps :rolleyes:

[riff_raff]

Because it's designed to be cheap, rather than efficient.I think what you meant to say is designed to be cheap to operate. Even an engine with high manufacturing cost can be very economical to operate if it is used many times and requires limited servicing.

You bring up a very interesting point regarding the trade off between cost and efficiency. Long before Elon Musk and SpaceX came along, there was a very bright entrepreneur named Andy Beal. Beal Aerospace's (http://www.freerepublic.com/focus/f-bloggers/1781167/posts) goal was building a heavy launch vehicle that would drastically reduce the $/lb cost of commercial launches. But rather than pursuing a reuseable design, his approach was to make the vehicle as simple and inexpensive to produce as possible. Of course the trade off was reduced performance.

The BA-2 heavy launcher (http://i6.tinypic.com/4ggtcwm.gif) used a single massive peroxide/kerosene engine for each of the first two stages. The engines were pressure fed from the propellant tanks and did not require turbo pumps. The composite propellant tanks were pressurized using helium. The engines used un-cooled (ablative) nozzles. Beal actually made quite a bit of progress with his limited resources. To give you an idea of the size of the engines, here's an image of the second stage engine being test fired:

http://hydrogen-peroxide.us/history-US-Beal-Aerospace/BA-2_stage-2-hot-fire.jpg

[joema]

When you consider the limited tools NASA had to work with, the shuttle was one of the greatest aerospace achievements in history. Besides all the issues with structures, dynamics and propulsion on ascent, no winged hypersonic vehicle (manned or unmanned) had ever reentered from orbit. There was very little data.

When the shuttle was designed in the mid-1970s, CFD was incredibly primitive, and hypersonic wind tunnels extremely limited. Yet the shuttle successfully flew a manned mission on the first flight.

Decades later, one reason (of many) NASP failed was continuing limitations in CFD and hypersonic wind tunnels. Unfortunately this is still a problem today, despite the much faster computers. This is due to limitations in CFD models and the fact that most demanding CFD problems are four dimensional. Every factor of two improvement in grid resolution requires a computational speed increase of 16 times. This is why hypersonic test vehicles like the X-43 and X-51 are so vital: even today, CFD cannot comprehensively simulate the required flow characteristics, and hypersonic wind tunnels are too limited in speed, duration and data quality.

This history is detailed in the book "Facing the Heat Barrier: A History of Hypersonics", which is available on line in PDF format:

Part 1: history.nasa.gov/sp4232-part1.pdf
Part 2: history.nasa.gov/sp4232-part2.pdf
Part 3: history.nasa.gov/sp4232-part3.pdf

SpaceX is doing a lot of leading edge work in CFD. Here is an interesting article with some links to some simulations:
Rockets Shake And Rattle, So SpaceX Rolls Homegrown CFD (http://www.nextplatform.com/2015/03/27/rockets-shake-and-rattle-so-spacex-rolls-homegrown-cfd/)

[riff_raff]

Indeed, one of the biggest engineering challenges facing the shuttle designers was aero compression heating on the structures. However, they did benefit from related work done on the X-15 program (http://history.nasa.gov/x15lect/structur.html).

Even SpaceX has wisely adopted technology developed by their competitors. They are using a material (Al-Li alloy) and construction method (friction stir welding) developed for the Delta IV common booster core, to reduce cost and improve reliability on their Falcon 9 vehicle.

Lastly, it was nice to see Blue Origin successfully land their booster a second time. They have had no problems with their landing gear, so maybe the mechanical systems engineering group at SpaceX could ask them for advice.:ok:

[Onceapilot]

Joema

A pleasure to read your posts... "Of course the ultimate goal is delivering a payload. On Apollo 15 the Saturn V delivered a payload of 140,930 kg (310,697 lbs) to low earth orbit. The heaviest shuttle payload was about 23,586 kg (52,000 lbs).

It would have taken six shuttle launches to deliver to LEO the same useful payload as a single Saturn V."

You are also correct to support LOX Rockets against air breathing space-plane lift to orbit, IMO. Cheers

OAP

[Dufo]

Shuttle payload was 23 tons but the Shuttle itself was additional 80 tons.

[riff_raff]

NASA was constantly working on propulsion system mods to increase the payload capacity of the shuttle. Some were put into service and some were only ground tested.
- The super lightweight (Al-Li alloy) external tank saved about 7,000lbs and was needed to get some of the heavier ISS module payloads to the required orbit.
- There was a lightweight filament wound SRB case developed but never flown.
- There was a 5 segment SRB developed for the shuttle that would have increased max weight of ISS payloads by about 10 tons. The 5 segment SRB is being used on the new SLS vehicle.
- I recall reading that NASA was looking at super cooling the LOx to reduce volume and allow a smaller, lighter tank to be used.
- The shuttle used the payload assist module (PAM) (https://upload.wikimedia.org/wikipedia/commons/0/04/SBS-3_with_PAM-D_stage.jpg) to boost payloads into higher orbits.

[GordonR_Cape]

A quick heads-up, this week is the 30th anniversary of the Challenger Shuttle disaster: https://en.wikipedia.org/wiki/Space_Shuttle_Challenger_disaster
Worth reading by anyone interested in the safety of complex technology.

[joema]

"A quick heads-up, this week is the 30th anniversary of the Challenger Shuttle disaster: https://en.wikipedia.org/wiki/Space_...enger_disaster
Worth reading by anyone interested in the safety of complex technology.
I wrote much of the original article, although I no longer maintain it.

National Geographic Channel is showing a documentary this week, titled "Challenger Disaster: Lost Tapes". It uses cinema verite techniques (no narrator, no new interviews, no commentary, no recreations) to factually explain the event using little-seen archival footage. Barbara Morgan (Christa McAuliffe's backup) described it as "very compelling and respectful".

Challenger Disaster: Lost Tapes - National Geographic Channel (http://www.natgeotv.com.au/tv/challenger-disaster-lost-tapes/)

About Challenger Disaster: Lost Tapes Show - National Geographic Channel - UK (http://natgeotv.com/uk/challenger-disaster-lost-tapes/about)

Link to trailer: http://tinyurl.com/zgkc7cx

For anyone interested in deep technical and procedural aspects, by far the most detailed account of the disaster and history of the SRB program is "Truth, Lies, and O-Rings: Inside the Space Shuttle Challenger Disaster", by Allan J. McDonald.

McDonald discusses what happened and implications for engineering ethics in this interview: https://www.youtube.com/watch?v=7_I-WUQvbjM

While a different incident, the transcript of the Columbia Accident Investigation Board (CAIB) from 4-23-03 is very educational, in particular statements by Robert F. Thompson, the shuttle program manager from inception to first flight. It covers key shuttle development decisions, development costs, planned flight rate, etc (do right-click and save as): https://www.nasa.gov/columbia/caib/PDFS/VOL6/H08.PDF If problems opening this, use top-level link and select "H.8 April 23,2003 Houston,Texas" : https://www.nasa.gov/columbia/caib/html/VOL6.html

[wiggy]

A quick heads-up, this week is the 30th anniversary of the Challenger Shuttle disaster

As an aside (i.e. thread drift) today is sadly the 49th Anniversary of the Apollo Fire.

http://history.nasa.gov/Apollo204/

[vaneyck]

While a different incident, the transcript of the Columbia Accident Investigation Board (CAIB) from 4-23-03 is very educational, in particular statements by Robert F. Thompson, the shuttle program manager from inception to first flight. It covers key shuttle development decisions, development costs, planned flight rate, etc (do right-click and save as): https://www.nasa.gov/columbia/caib/PDFS/VOL6/H08.PDF If problems opening this, use top-level link and select "H.8 April 23,2003 Houston,Texas" : https://www.nasa.gov/columbia/caib/html/VOL6.html
Thank you very much for this, joema, it's fascinating. As it happens I just got through re-reading the main CAIB final report - one of the best-written and organised reports I've ever come across - but I had none of the later appendices (Appendices A, B, and C are included in the final report). You've provided me with some more good reading.