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 The propulsion systems available and capable of performing an abort during this phase with the irrespective $$\Delta\mathrm{V}$$ capa­bilities are shown in Table I. Note that the ServiceModule RCS propuls ion system does not appear on this table The reason for this is that when considering only abort trajectories targeted to the center of the reentry corridor, the Service Module RCS does not have sufficient $$\Delta\mathrm{V}$$ capa­bility to perform an abort maneuver from the nominally planned earth parking orbit. However, a procedure is currently being evaluated where the abort maneuver is targeted for very near the over shoot boundary of the reentry corridor. Preliminary indications are that an abort using this technique will be available using the Service Module RCS propulsion system, although it will be marginal. As shown in Table I, the propulsion systems which are available and capable of performing an abort are the SPS, the LM propulsion systems, and the S-IVB.

Figure 6 summarizes the abort modes for earth parking orbit. The first and primary mode of interest, as shown, is a single coplanar deorbit burn targeted to provide horizon monitoring and which results in landing at a discrete area. The initia­tion time of this type of abort is carefully selected to pro­vide the landing area control. This mode is, of course, very similar to that planned for Project's Mercury and Gemini. As shown in the chart of figure 6, the abort can be performed by either the SPS or S-IVB throughout the entire phase with, of course, the SPS being the prime propulsion system. Use of the S-IVB is not desirable due to possible recontact problems during reentry and, thus, would never be considered as an abort mode unless there had been a definite indication by the instrumentation that an SPS failure had occurred prior to CSM separation from the S-IVB. Also, if use of the Service Module RCS to deorbit proves feasible by targeting reentry near the overshoot boundary, the possible use of the S-IVB as an abort propulsion system would seem even more remote. Thus, use of the S-IVB as an abort propulsion system during this phase, seems very improbable even though it is available.

Figure 7 shows the major features of the SPS and S-IVB abort mode during this phase. Note that the transfer angle from the abort maneuver point to reentry is much less than 90°, which means the time from abort to reentry is on the order of 15 to 20 minutes. Command Module/Service Module separa­tion occurs during the coast period from abort to re entry followed by the Command Module orienting itself for reentry. The burn attitude is such that the maneuver is coplanar and such that the earth's horizon remains at a fixed position in the Command Module window for crew monitoring purposes. The $$\Delta\mathrm{V}$$ required for abort is approximately 500 fps, which is, of course, well within the SPS capability. The time of de- ­ orbit is selected so as to cause landing in a discrete recovery area, as mentioned previously.