See list attachedOctober 25, 196868-PA-T-238APA/Chief, Apollo Data Priority CoordinationDescent Aborts – Part II
This memo is to carry on from that three page snowflake I sent you the other day on the same subject. It turns out we have encountered one of those rare situations when in doing something to fix an undesirable situation we actually improve something else at the same time. Speci- fically, the rendezvous people want to target the LM to a substantially higher orbit following an early descent abort than they had previously proposed. This makes the horizontal posigrade burn following the descent abort larger, of course, and alleviates that crazy pitch profile problem which used to exist during an abort in the first 50 seconds of powered descent. The point is that by some fairly minor changes in the space- craft computer program (LUMINARY), we can probably eliminate the special abort procedure we used to think was necessary early in descent. Changes to the DPS abort program (P70) are essentially just changes in some erasible constants. This does not impact coding but has a significant impact on testing. By that, I mean the program will work now. The APS program change noted in last week's memo is still required but is essentially achieved by a erasible constant change too. This will all be firmed up and brought to the Software Configuration Control Board in the near future for their approval or something.
Having the early abort situation under control, we pressed on to another phase of descent aborts requiring some attention – specifically, how to handle the situation when the DPS is not quite capable of getting the LM all the way back into the desired insertion orbit. In order to establish procedures, it was necessary to make some assumptions. They are:
1. We never want to “Abort Stage” and use the APS, if the DPS is still operational.
2. It is acceptable to operate the DPS to propellant depletion.*
3. We have no desire to use the APS engine again after achieving orbit (that is, during rendezvous). Of course, we intend to use the APS propellant through the RCS interconnect.
* This assumption must be verified by ASPO and then included in their data books.
4. The “Abort Monitor” in LUMINARY remains active following a DPS propellant depletion cutoff, which may result in a ΔV monitor alarm, even though the crew calls up the ΔV residuals.*
If we can make the above assumptions, the procedures become quite simple and standard. Namely, whenever aborting on DPS, the crew will permit that engine to operate at full thrust until either a guided cutoff is achieved or propellant depletion occurs. At that time, the crew will “proceed” to the DSKY display of ΔV residuals. If the ΔV remaining to be gained is less than 30 fps, the DPS will be manually staged and the crew will utilize the RCS to achieve the desired insertion condition by nulling the ΔV residuals. (It is probable that only the horizontal component need be trimmed if a convenient attitude reference is available. The FDAI eight ball should be good for this.) If the ΔV to be gained is in excess of 30 fps, the crew will hit “Abort Stage,” automatically jettisoning the DPS and lighting off the APS to make up the ΔV deficiency. Again, only the horizontal ΔV residual need be trimmed.
It is to be noted that with the new, high apogee we will be targeting for, the RCS/APS switchover point is orbital by a substantial margin (apogee in excess of 75 miles) and so there is no problem in the use of an RCS burn whose duration is less than 30 seconds. It is also to be noted that if the ΔV required of the APS is less than 100 fps, the burn duration will be less than 10 seconds, which probably makes it unsafe to reignite the APS. There is so much mystery with what is and what is not acceptable with the APS we cannot really be sure about that. However, it does not matter since there is no problem anticipated in performing the rest of the maneuvers with RCS.
One final comment – it has been proposed that the DPS be operated at half thrust during aborts to prevent lofting when the APS is required to achieve orbit. Two miles perigee and four miles apogee are the maximum effects. Those do not significantly perturb the abort rendezvous and therefore the decision was to maintain full thrust.
* This assumption must be verified by me with MIT.