Questions about Saturn V and Apollo

[Errol]

Doing some research for an Apollo recreation I'd like to do, I'm having trouble finding answers to some questions I have about the fuel systems and engines. Realizing that each mission had different amounts of payload, and that the rocket design was changed to facilitate more payload for later missions (such as reducing the number of ullage and retro rockets used for staging throughout the first three stages), lets just pick Apollo 11 to base all these numbers off of.
 

First major question I have is the thrust to weight ratio for each stage. I've read at liftoff the rocket had a TWR of about 1.2, and that the ascent stage of the LEM had a TWR of about 2.1 at liftoff in lunar gravity. I'm not sure how accurate those are (I don't trust wikipedia), and I'm lacking any information about the other stages. I've been able to find references to the burn times for each stage fairly easily, but I can't seem to pin down the thrust to weight ratios. So to list what I am missing, I need the TWR for S-II, SIV-B, the CSM, LEM Descent stage (lunar TWR, to be clear) and also for the Launch Escape System (with the CM attached).
 

The other major area of concern I have that I can't quite figure out is the fuel systems routing and ullage requirements for the LEM. I've read that the same hypergolic fuel was used for the descent engine, ascent engine, AND the RCS thrusters, is this correct? I've seen diagrams that show elliptical fuel tanks (with their helium pressurization tanks as well) in the descent stage that I believe require ullage burns from the RCS quads on the ascent stage. I've also seen diagrams of the ascent stage that show both elliptical and spherical fuel tanks, but no indication of if the ascent motor requires ullage burns from the RCS before firing. I have several questions/gaps in my understanding here. Was the RCS only supplied by the pressurized spherical tanks in the ascent stage, or was there some way for them to draw fuel from the descent stage so that when they launched the ascent stage it would have full tanks for RCS?
 

Last question is more broadly speaking about all the missions that went to the moon at all. It's one detail about the flight plan that I seem to have found evidence for two different possibilities. I'm wondering about the trans-lunar injection burn. Was the burn one long continuous burn from stage one on the ground, all the way to the moon, with the correct time of day used to ensure the correct phase angle for departure from earth OR did they cut the engines after circularizing in LEO, and then re-light the SIV-B for the Hohmann transfer?

[StrandedonEarth]

I do know that Saturn V Moon missions launched into a parking orbit where system checkouts were performed, before relighting the third stage for TLI. As for the rest, no idea

Edited January 3 by StrandedonEarth

[sevenperforce]

4 hours ago, Errol said:

First major question I have is the thrust to weight ratio for each stage. [snip] I've been able to find references to the burn times for each stage fairly easily, but I can't seem to pin down the thrust to weight ratios. So to list what I am missing, I need the TWR for S-II, SIV-B, the CSM, LEM Descent stage (lunar TWR, to be clear) and also for the Launch Escape System (with the CM attached).

To answer these questions reliably, your best approach is to look at the actual thrust of each engine, the number of engines on each stage, and the weight of the stack at each staging event. Then just do the math. All of that is going to be much more accurate than guessing at whether publicly-posted numbers were posted by people who did the math correctly. The wet and dry masses of each stage is all public from NASA documents.

4 hours ago, Errol said:

The other major area of concern I have that I can't quite figure out is the fuel systems routing and ullage requirements for the LEM.

I happen to know a good bit about this so you're in luck!

4 hours ago, Errol said:

I've read that the same hypergolic fuel was used for the descent engine, ascent engine, AND the RCS thrusters, is this correct?

The same fuel type was used, yes, but it was not sourced from the same tanks.

The RCS thrusters on the ascent stage provided 100% of the reaction control for both the descent and ascent; there were no RCS thrusters on the descent stage. The tanks on the descent and ascent stages were completely separate without any interconnections. The descent propulsion engine was a throttleable, restartable, gimballed hypergolic engine fed exclusively from tanks housed within the descent stage. Because this engine could be gimballed, the RCS on the ascent stage was used only for roll control while the descent stage was firing and ullage while the descent stage was starting up. While I know that the RCS thruster controls were designed to permit translational burns for docking, I am not sure whether translational firing was active during hover and landing. I should also note that although the lunar module was capable of acting as the active translational actor during docking, it never did so in practice; it would just hold orientation and allow the CSM to come to it.

The ascent stage had three sets of propellant tanks: one main set of propellant tanks which fed the ascent propulsion system and a pair of redundant propellant tank sets for the RCS system. Each of the redundant RCS propellant tank systems contained half the propellant needed for descent RCS plus all of the propellant needed for ascent RCS, so that if one of the tanks experienced a problem they could still perform all necessary ascent burns. There was an additional (closed) valve linking the main propulsion propellant tanks to the RCS system so that the RCS system could be powered directly from the main prop tanks if both sets of RCS system tanks failed, although this would reduce the amount of propellant available for the ascent propulsion system.

The ascent propulsion system was constant-thrust and fixed, so the RCS system had to provide pitch and yaw as well as roll during ascent. Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts. There was no ullage burn off the lunar surface because lunar gravity provided sufficient propellant settling.

4 hours ago, Errol said:

Last question is more broadly speaking about all the missions that went to the moon at all. It's one detail about the flight plan that I seem to have found evidence for two different possibilities. I'm wondering about the trans-lunar injection burn. Was the burn one long continuous burn from stage one on the ground, all the way to the moon, with the correct time of day used to ensure the correct phase angle for departure from earth OR did they cut the engines after circularizing in LEO, and then re-light the SIV-B for the Hohmann transfer?

@StrandedonEarth is correct: the SIV-B did a single burn from staging to parking orbit, circled Earth a few times, then restarted for the trans-lunar injection burn. The parking orbit was both to simplify phasing and to allow time for systems check-out before committing to the rest of the mission. Oberth, drag, and boiloff losses were accounted for but minimal.

[Errol]

First, thank you for the detailed reply.

23 hours ago, sevenperforce said:

To answer these questions reliably, your best approach is to look at the actual thrust of each engine, the number of engines on each stage, and the weight of the stack at each staging event. Then just do the math. All of that is going to be much more accurate than guessing at whether publicly-posted numbers were posted by people who did the math correctly. The wet and dry masses of each stage is all public from NASA documents.

Initially I had tried this (not by hand, my math is shaky at best, so I used a TWR calculator script I found online). I noticed my numbers for the LEM descent stage were way out of wack. Then I realized I wasn't sure if you are supposed to compensate for lunar gravity by adjusting just the weight or both the thrust and the weight. Anyway, at this point with my level of uncertainty I decided to just add it to the list of things to ask about in this thread.
 

23 hours ago, sevenperforce said:

I should also note that although the lunar module was capable of acting as the active translational actor during docking, it never did so in practice; it would just hold orientation and allow the CSM to come to it.

I had heard about the ascent module not ever being the active translational actor during docking, but wasn't the ascent stage responsible for all of the phasing/catch-up/rendezvous maneuvers, only switching to station keeping once close enough for final approach to be initiated by the CSM?
 

23 hours ago, sevenperforce said:

 There was an additional (closed) valve linking the main propulsion propellant tanks to the RCS system so that the RCS system could be powered directly from the main prop tanks if both sets of RCS system tanks failed, although this would reduce the amount of propellant available for the ascent propulsion system.

Do you know if there were any other procedures required if this valve was going to be used? Was there some sort of additional pressurization hardware (like a bladder or something) to enable this, otherwise how would the main tanks deal with ullage?
 

23 hours ago, sevenperforce said:

Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts.

What was tricky about it? Also, if the ascent stage was responsible for rendezvous maneuvers like I asked about above, does this mean that they only used the RCS translation for those burns?
 

23 hours ago, sevenperforce said:

@StrandedonEarth is correct: the SIV-B did a single burn from staging to parking orbit, circled Earth a few times, then restarted for the trans-lunar injection burn. The parking orbit was both to simplify phasing and to allow time for systems check-out before committing to the rest of the mission. Oberth, drag, and boiloff losses were accounted for but minimal.

 

Does this mean that the CSM RCS was used for the TLI ullage burn? EDIT: I've just learned about the S-IVB auxiliary propulsion system....essentially an RCS system running on pressurized hypergolic propellant. It was used for roll, pitch and yaw, as well as ullage.

Edited January 4 by Errol

[sevenperforce]

14 minutes ago, Errol said:

First, thank you for the detailed reply.

No problem.

14 minutes ago, Errol said:

Initially I had tried this (not by hand, my math is shaky at best, so I used a TWR calculator script I found online). I wasn't sure if you are supposed to compensate for lunar gravity by adjusting just the weight or both the thrust and the weight.

I would suggest making a spreadsheet, carefully laying out the dry mass and propellant capacity of each stage, and so forth.

How you compensate for lunar gravity is up to you; just keep it consistent. The TWR is the total thrust of all firing engines divided by the weight of the vehicle; the weight of the vehicle is its mass times the gravitational acceleration. Personally, I just imagine it's Earth gravity for everything and then if I am dealing with lunar touchdown or takeoff I apply the appropriate transformation afterward. 

14 minutes ago, Errol said:

I had heard about the ascent module not ever being the active translational actor during docking, but wasn't the ascent stage responsible for all of the phasing/catch-up/rendezvous maneuvers, only switching to station keeping once close enough for final approach to be initiated by the CSM?

Yes, that's correct.

The lunar module ascent had two phases: a ten-second burn straight up to clear terrain, and then a pitchover and burn to orbit. The precise moment and orientation of liftoff was chosen to minimize phasing and plane changes. If the ascent propulsion engine had failed during the last thirty seconds of the orbital insertion, the four little aft-facing RCS thrusters had sufficient umph to complete orbital insertion. That was one of the reasons for the valve that would allow the main tank propellants to flow directly into the RCS thrusters.

14 minutes ago, Errol said:

[snip]

Also, if the ascent stage was responsible for rendezvous maneuvers like I asked about above, does this mean that they only used the RCS translation for those burns?

Yes; following successful orbital insertion, the rendezvous maneuvers were performed entirely by the RCS system. Although the RCS system had three-plane translational capability, all of these maneuvers were being done pretty much manually, often using slide rules to compute the proper heading, time, and duration of each burn. As a result, the aft-facing RCS thrusters were the only ones used for these maneuvers; the other thrusters only provided attitude control.

For Apollo 11, the initial launch reached an 87.6x17.6 km elliptical orbit at main propulsion system burnout. RCS was used about an hour later, at apoapsis, to circularize. They had planned for a plane change burn as well but didn't need it. By this point, they were nearly on a collision course with the CSM so they did a Constant Delta Height burn (still with RCS) which ensured that the two vehicles would be constantly 28 km apart and they would have time to plan the remaining rendezvous burns. 

As an example -- the RCS circulation burn for Apollo 11 required 15.7 m/s of dV or around 13.7 kg of propellant. The burn took just over two minutes, starting at MET 125:19:34.70 and ending at around MET 125:21:36. Had the burn been performed with the actual ascent propulsion engine, it would have taken less than three seconds.

14 minutes ago, Errol said:

Do you know if there were any other procedures required if this valve was going to be used? Was there some sort of additional pressurization hardware (like a bladder or something) to enable this, otherwise how would the main tanks deal with ullage?

I don't know the specific procedures but it definitely all would have happened pretty quickly, so it was likely automated.

The RCS propellant tanks used teflon bladders to hold the propellant inside a pressurized tank. As helium pressurant was vented into the tank, the teflon bladders were compressed and pushed the RCS propellant out, obviating the need for any separate RCS ullage burn. If all of the RCS tanks had COMPLETELY failed and the main tank was being used exclusively, then ullage burns would become a problem. There was a chance that the surface tension of the propellant in the feed lines would be enough for the initial ullage puff but it would have been tricky. But that was an unlikely contingency (lots of other stuff would have to fail which would probably be LOCV anyway).

14 minutes ago, Errol said:

On 1/3/2024 at 5:08 PM, sevenperforce said:

Although the ascent propulsion system was technically restartable, it was not ordinarily restarted during ascents because the helium pressurization for restarts was a little tricky: it was a single burn from the lunar surface to lunar orbit. It was restarted for disposal burns, and required an RCS ullage burn for those restarts.

What was tricky about it?

It used multiple helium tanks with burst discs (for simplicity) so it could only be started twice.

14 minutes ago, Errol said:

Does this mean that the CSM RCS was used for the TLI ullage burn?

No, definitely not. The CSM never did anything until separation from the third stage.

The S-IVB third stage of the Saturn V was equipped with a pair of Auxiliary Propulsion System modules using hypergolic propellants. Each module carried around 120 kg of propellant and boasted a trio of RCS engines and a single ullage engine. They were used to provide ullage for third-stage restarts, roll control (and backup pitch/yaw control) during the third stage burns, and general attitude control during the transposition and docking maneuver. 

[kerbiloid]

Just is looking good and on topic.

https://history.nasa.gov/SP-4029/Apollo_11i_Timeline.htm

https://www.c82.net/work/?id=368