The second & third paragraphs reference Phantom Wolf's work in demonstrating how the behavior of the Earth's image invalidates Sibrel's hypothesis. Depending on how you reference his work, you may want to tweak the text to accomodate your format.
I left it as you originally wrote it, except I referred to PhantomWolf as an “Apollohoax.com member” to explain the use of an alias.
PhantomWolf, do you prefer I use your real name?
If you’ve got any other edits you want to make, please speak up.
Joined: May 2005 Gender: Male Posts: 4,620 Location: Lost Deimos Moonbase
Re: Bob Braeunig's Apollo Pages « Reply #17 on Dec 6, 2009, 4:35pm »
To be honest, bob, few people in the debate know me by my real name, lol. About the only place I ever use it is on IMDB, which I post very infrequently on, and that is because they use emails as users so I am stephen_a_h. PW is fine. I'm also planning to update my own page (including a movie of the moon moving behind the window rather than just the still images, tiding up the text and giving better information, and adding some "new" stuff to it that gives even more evidence, including images taken from weather satellites on the day.)
It must be fun to lead a life completely unburdened by reality. -- JayUtah
"On two occasions, I have been asked, 'Pray, Mr. Babbage, if you put into the machine wrong figures, will the right answers come out?' I am not able to rightly apprehend the kind of confusion of ideas that could provoke such a question." -- Charles Babbage (1791-1871)
Joined: May 2005 Gender: Male Posts: 3,072 Location: Dayton, Ohio
Re: Bob Braeunig's Apollo Pages « Reply #22 on Jul 26, 2010, 8:23pm »
After shelving the project for several months, I finally got around to finishing my Saturn V launch simulation. I got a pretty good match to actual ascent data but I had to take a few liberties to get it to work, so I’m not 100% satisfied. If anybody cares, here’s the link:
Joined: Jun 2005 Posts: 1,119 Location: I come from a land down under
Re: Bob Braeunig's Apollo Pages « Reply #23 on Jul 26, 2010, 9:32pm »
Bob
Impressive piece of work. *golf clap*
Quite a few years ago I did a very rough calculation for the acceleration of the first stage only on a spreadsheet. I only looked at altitude, so the result was only indicative. But all of a sudden I understood why they shut down the centre engine 15 seconds ahead of the other four - I simply hadn't realised how much of the Saturn V's mass was in the first stage fuel tanks.
However, I remain puzzled by the PU shift for the second stage. The Apollo Flight Journal says it was "...part of a strategy to make sure that as little propellant as possible is left in the tanks when the second stage has done its work...." I just don't understand how that's so. Why not just put a bit less LOH into the tank in the first place and leave the fuel/oxidiser mix unchanged?
Incidentally, the AFJ specifies different PU ratios than the ones you use.
I simply hadn't realised how much of the Saturn V's mass was in the first stage fuel tanks.
Oh yes. The purpose of that inboard shutdown is to limit acceleration to 4 g. Check out the flight report for a typical Saturn V flight. They always include a plot of longitudinal acceleration vs time. They're very characteristic, so if you've seen one they're very easy to find in the other reports.
I wonder how much theoretical performance they lose by doing that. Gravity loss is always greatest early in the flight of a launcher, when the pitch angle is still high and a significant fraction of thrust is still overcoming gravity.
That's why the Saturn V's acceleration is greatest during first stage flight (and why kerosene was chosen as the fuel.) The upper stages operate with a pitch angle of nearly zero, so the gravity loss is correspondingly small. Then the engines can be optimized for Isp rather than maximum thrust.
Quote:
However, I remain puzzled by the PU shift for the second stage. The Apollo Flight Journal says it was "...part of a strategy to make sure that as little propellant as possible is left in the tanks when the second stage has done its work...." I just don't understand how that's so. Why not just put a bit less LOH into the tank in the first place and leave the fuel/oxidiser mix unchanged
The problem is that you don't always know, preflight, exactly what the fuel-oxidizer flow rates will be for those particular engines under those particular conditions. The IU can measure them in real time and adjust the time of PU shift so that the two propellants will deplete at (or close to) the same time.
Also note that when the PU shift occurs and the mixture goes a little rich, the Isp of the engines increases slightly due to the decrease in average molecular weight of the exhaust. I don't know how much this contributes to performance - it may not be significant - but it's still there. You don't want to do it earlier because the PU shift also decreases thrust, and decreasing thrust before you've pitched down to the local horizontal increases gravity loss.
However, I remain puzzled by the PU shift for the second stage. The Apollo Flight Journal says it was "...part of a strategy to make sure that as little propellant as possible is left in the tanks when the second stage has done its work...." I just don't understand how that's so. Why not just put a bit less LOH into the tank in the first place and leave the fuel/oxidiser mix unchanged?
ka9q has already explained this well, there's not much I can add.
They always include a plot of longitudinal acceleration vs time. They're very characteristic, so if you've seen one they're very easy to find in the other reports.
An acceleration vs. time graph is included with my simulation.
I wonder how much theoretical performance they lose by doing that. Gravity loss is always greatest early in the flight of a launcher, when the pitch angle is still high and a significant fraction of thrust is still overcoming gravity.
Center engine cutoff occurred with the flight path angle about 23 degrees, so the rocket had already transitioned to mostly horizontal flight. I haven’t simulated it, but I doubt much performance was lost due to center engine cutoff.
Also note that when the PU shift occurs and the mixture goes a little rich, the Isp of the engines increases slightly due to the decrease in average molecular weight of the exhaust.
That’s true of the J-2 engine because it used LOX/LH2. Most other engines operate with their mixtures ratios at, or very near to, that which maximizes specific impulse. Shifting the MR, therefore, would lower ISP.
LOX/LH2 engines are different in that maximizing ISP is not the only major consideration when choosing MR. Maximum ISP may be achieved with the MR somewhere around 3.5. The problem is that LH2 has an extremely low density and requires very large storage tanks. Setting the MR as low as 3.5 means the rocket gets very large, resulting in greater tank mass and more drag. Increasing the MR to 5 or 6 means the engine operates at a slightly lower ISP, but the average propellant density increases significantly. The advantage of lower propellant volume more than offsets the small decrease in ISP.
I don't know how much this contributes to performance - it may not be significant - but it's still there.
The published data that I’ve seen says the J-2 has an ISP of 424 s at MR 5.5 and 427 s at MR 5.0. I ran some numbers and estimated an ISP of about 434 s at MR 4.5.
For comparison, the Apollo 11 data that I based my simulation on gives the following for the first burn of the S-IVB:
Re: Bob Braeunig's Apollo Pages « Reply #27 on Jul 27, 2010, 6:03pm »
Good points about extremely low density of LH2, the fuel for the J-2s on the S-II and S-IVB. And you're right that not every upper stage burns LH2. A rich mixture is more likely to benefit Isp in a LH2-burning engine because of the low molecular weight of LH2.
Hypergolic upper stages are still fairly common, especially when a coast phase is required because they're especially easy to restart. But the general trend seems to be toward LH2-burning upper stages because of their nearly unbeatable Isps. (What's better? LH2 + LF2 - liquid fluorine? I hope nobody ever flies that.) The Falcon is a singular (?) recent exception that burns kerosine in its upper stage.
I've seen the tank size/weight argument given for why the S-IC burned kerosene rather than LH2. But what about thrust/weight? During first stage flight the Saturn V achieved 4 g (and had to be limited there by shutting down an engine) despite being considerably more massive than the upper stages, which only achieved about 2 g, max. The kerosene-burning F1s seemed to be optimized for thrust while the LH2-burning J2s were optimized for Isp, and that's as it should be since gravity losses are most significant during first stage flight.
You see the same principle in the widespread use of solid rocket boosters during first stage flight. Solids can achieve high thrust/weight ratios more easily than liquid fueled rockets, though their Isps are again limited.
I believe the SPS on the Apollo CSM also had a propellant utilization valve. Because that engine was hypergolic, I suppose it was continually adjusted to maintain a single, fixed and optimum fuel/oxidizer ratio.
I don't recall reading about any PU valves on the LM.
After shelving the project for several months, I finally got around to finishing my Saturn V launch simulation.
Very nice work, Bob!
You wonder why the propellant flow rate increased during the S-IC and S-II flights, and speculate that it was caused by the increase in longitudinal acceleration increasing the pump inlet pressures.
I think this is almost certainly the case. The trouble both stages had with pogo lends further weight to this theory - pogo happens when the propellant flow rate (and thrust) increases with acceleration even when you don't want it to.
The "fix" for pogo on the S-IC involved adding a helium pressurized "accumulator" to the propellant line. This would have dampened out rapid changes in inlet pressure due to the inertia of the propellant in the lines, but I don't think it would have changed the average flow rate vs acceleration. So the effect would still be there.
As I recall, pogo on the S-II was never really solved; it was simply avoided by shutting down the inboard engine early (yet it still happened on Apollo 13). By limiting peak acceleration as the stage lost mass, that would have reduced the magnitude of the effect on thrust. But it would still be there.
Maybe it's worth comparing the pump inlet pressures to the pressures maintained in the tank ullages vs the pressure due to the (accelerated) weights of the propellants. Remember that the head of propellant in each tank decreases as it empties.
You wonder why the propellant flow rate increased during the S-IC and S-II flights, and speculate that it was caused by the increase in longitudinal acceleration increasing the pump inlet pressures.
I think this is almost certainly the case.
Thanks, I'm happy to get a second opinion on that.
Maybe it's worth comparing the pump inlet pressures to the pressures maintained in the tank ullages vs the pressure due to the (accelerated) weights of the propellants. Remember that the head of propellant in each tank decreases as it empties.
The LOX tank ullage was kept pressurized during flight to between 18 and 23 psia by feeding GOX from the F-1 heat exchanger to the top of the tank. The length of the LOX tank was 64 feet and the bottom of the tank was about 55 feet above the top of the engines (length of the stage – length of LOX tank – length of engine = 138’ – 64’ – 19’ = 55’).
Let’s take the points when the tank was 2/3 full and 1/3 full. I estimate the liquid levels to be about 41’ and 23’ respectively, giving us heads of 96’ and 78’. Using my simulation, I estimate these points occur at 51.5 s and 105 s after launch, at which time the accelerations are 1.67 g to 2.61 g.
Let's use the average ullage pressure, 20.5 psia, and add the liquid head to get the pressure at the pump suction. The density of LOX is 71 lbm/ft3, therefore we have:
@ 2/3 full: 20.5 psia + 96 ft x 71 lb/ft3 x 1.67 / 144 in2/ft2 = 100 psia
@ 1/3 full: 20.5 psia + 78 ft x 71 lb/ft3 x 2.61 / 144 in2/ft2 = 121 psia
The fuel tank was also kept pressurized during flight with gaseous helium, but I can't find the pressure. Let's assume the same as the LOX tank. The length of the fuel tank was 43 feet and the bottom of the tank was about 10 feet above the top of the engine.
At the points when the tank was 2/3 full and 1/3 full, I estimate the liquid levels to be about 27’ and 16’ respectively, giving us heads of 37’ and 26’. The density of RP-1 is 50 lbm/ft3, therefore we have:
@ 2/3 full: 20.5 psia + 37 ft x 50 lb/ft3 x 1.67 / 144 in2/ft2 = 42 psia
@ 1/3 full: 20.5 psia + 26 ft x 50 lb/ft3 x 2.61 / 144 in2/ft2 = 44 psia
So it looks like the pressure change at the fuel pump suction is small, but a 21 psia change at the oxidizer pump suction should be enough to produce a noticeable change.
Please correct me if I'm wrong (it's a long time since I last dealt with problems involving pumps), but doesn't a pump simply add pressure? That is, increasing the pressure on the suction side should produce an equivalent increase in pressure on the discharge side. Higher discharge pressure means we have a greater pressure differential across the engine injector, thus an increase in flow rate.
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