Status of the lunar module "quick return" guidance capability
- Date: Sep 2 1966
- From: FM/Deputy Chief, Mission Planning and Analysis Division
- Subject: Status of the lunar module "quick return" guidance capability
This note is in response to your query regarding the “quick return” capability being provided in the lunar module (LM) for aborts during the lunar descent phase. As you recall, I reported deletion of a program in the LM computer for generating coefficients to be used in an abort polynomial to retarget the LM powered flight to provide a direct intercept rendezvous trajectory. You asked how far this work had progressed since you felt such a capability would be “comfortable.”
In answer to that question, MIT informed me that, whereas the concepts were well established, there was still a considerable amount of work required to complete this particular program. Furthermore, we have also deleted the direct ascent launch guidance, which is a necessary companion program. Certainly of more interest to you now is, what is our current capability.
The program is being written such that abort action by the pilot during powered descent will cause the guidance to retarget to the standard LM insertion orbit. Incidentally, it is necessary for the astronaut to select which engine, the Ascent Propulsion System (APS) or the Descent Propulsion System (DPS), is to be used, depending on the situation. In any case, following insertion into orbit, the crew has two choices: either to proceed with the concentric flight plan, or to use a processor which we have retained for just such situations as this, whereby the crew may obtain the two-impulse Lambert solution for rendezvousing with minimum ΔV–essentially a direct intercept. In effect, the latter provides very nearly the same capability as we have deleted, except that the maneuver must be carried out in two steps with some delay–say, five or ten minutes–between them, as opposed to a single maneuver.
If the concentric flight plan is chosen, the time between the abort action and rendezvous would be about 2½ hours with the differential altitude varying between 42 nautical miles above to the standard 15 nautical miles below the CSM, depending on whether the abort took place immediately after initiation of the descent maneuver or at the end of the hover. The “direct intercept” approach would take about l½ hours but is only possible prior to initiation of hover since after that time the intercept trajectory, unfortunately, also intercepts the moon–first! Actual procedures have to be settled, but I feel we’re in pretty good shape here.
Finally, regarding current status, there are some unresolved problems associated with this retargetting which MIT is currently addressing. For example, if an abort occurs early in descent, the LM will be near 50,000 feet with orbital velocity. The current insertion altitude is 60,000 feet. Thus, the spacecraft would have to make a large altitude change with little increase in velocity, which would obviously demand some rather wild gyrations of the spacecraft–both highly undesirable and unnecessary.