# Apollo Lunar Landing Mission Symposium/Apollo Earth Return Abort Capabilities

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APOLLO EARTH RETURN ABORT CAPABILITIES
By
Ronald L. Berry
APOLLO EARTH RETURN ABORT CAPABILITIES
It has always been a stated desire, as regards Project Apollo, to have continuous abort-to-earth capability throughout the entire lunar landing mission. This paper will examine the capability of the spacecraft to satisfy this objective through each mission phase of the lunar landing mission. The abort capability will be discussed primarily from a performance standpoint. In other words examining whether or not the spacecraft has the necessary performance required for continuous abort capability through— out the mission, Portions of the mission where redundancy exists will be pointed out as well as portions of the mis- sion which are critical or marginal as regards abort capa- bility. The primary ground rules for this discussion are as follows;
The items which will be discussed for each mission phase are as follows:
The basic abort modes or procedures will also be briefly described. Where possible, the mode considered prime will be pointed out along with the modes which are optional or backup, The computer logic required to provide all of these abort modes and procedures are currently
The capability of the spacecraft of actually performing abort maneuvers will be disscussed. Included here will ​be a descrption or the propulsion systems available, the $\Delta \mathrm {V}$ available from each, and the $\Delta \mathrm {V}$ required for abort.
$\Delta \mathrm {V}$
The major constraints which shape all abort trajectories are as follows:
$\Delta \mathrm {V}$
$\Delta \mathrm {V}$
$\Delta \mathrm {V}$
$\Delta \mathrm {V}$
The mission phases considered in this discussion will be (1) launch—to—earth parking orbit, (2) earth parking orbit coast, (3) translunar injection, (4) translunar coast, (5) lunar orbit insertion, (6) lunar orbit cost (7) transearth injection, and (8) transearth coast. Note that aborts during LM maneuvers are not considered (illegible text) would be to rendezvous.
ABORTS DURING LAUNCH PHASE
(illegible text) ​The propulsion systems available for abort during the launch phase are as follows:
Figure 1 shows a summary of the abort modes for the launch phase and through what region of the launch burn they apply. These modes are shown as a function of the burn time as well as the respective launch vehicle stage. The black bar des- ­ ignates the prime mode while the striped bar represents the optional or backup modes of abort.
Note that the first mode is the LES or launch escape system mode which is prime from the pad throughout the S-I stage and extending on for a few seconds into the S-II burn before LES, jettison. As shown in figure 2, a LES abort consists of the LES propulsion system separating the Commnand Module from the stacked launch vehicle configuration and providing an adequate altitude and downrange translation. This is followed by the orientation of the Command Module with heat shield for- ­ ward for reentry. Landing occurs in a continuous Atlantic recovery area along the flight azimuth up to a maximum down­ range of approximately 400 nautical miles.
The next mode is the suborbital free-fall abort. This mode, as shown in the summary chart, begins where the LES is jettsoned and remains available until approximately halfway through the S-IVB burn. Note that this mode is prime through approximately the first half of the S-II burn and through half of the S-IVB burn. It is considered an optional mode through the second half of the S-II burn because of the availability of a contingency; orbit insertion with the S-IVB, which will be discussed below.
As shown in figure 3, the suborbital free-fall mode consists of CSM separation from the launch vehicle using Service Module RCS, a 10-second SPS burn to gain further separation from the launch vehicle, Service Module jettison, and Command Module orientation (heat shield forward) for re entry. Landing would be in the continuous Atlantic recovery area along the flight azimuth up to 3,200 nautical miles downrange.
An extension of the suborbital abort mode can be achieved by addition of another SPS burn for landing area control, as shown in the summary of abort modes in figure 1. This mode is available as an option during the second half of the S-IVB burn when the suborbital free-fall mode is no longer available.
​Figure 4 shows how the suborbital mode with SPS landing control differs from the free-fall mode. Note that the procedure is identical except for an additional retrograde SPS burn.
This burn is a variable length depending on the time of abort and causes landing to be at the end of the continuous Atlantic recovery area, approximately 3,200 nautical miles downrange. The next mode of interest, as shown in figure 1, is the S-IVB contingency orbit insertion followed by an SPS de-orbit to reentry. This mode, as shown, is available and prime for approximately the second half of the S-II burn. The reason that this mode is considered prime over the suborbital free-fall mode is that it allows landing to be pinpointed precisely to a given recovery force and would allow additional "thinking" time to consider alternate missions.
The next mode is similar in nature to the S-IVB contingency orbit insertion, except the SPS is used for both the COI and the deorb it burn.
This mode, as shown in figure 1, is only possible during approximately the latter half of the S-IVB burn, but it is the prime mode for this time period.
Figure 5 shows the basic features of these latter two abort modes.
As shown, insertion into earth parking orbit is com­pleted by e ither the S-IVB or the SPS. After a certain coast time in earth parking orbit, during which CSM S-IVB separation occurs (if it has not already), an SPS coplanar deorbit burn is performed to return the spacecraft to reentry. The time of deorbit is chosen so as to result in landing at a discrete recovery area, as done in Projects Mercury and Gemini. The SPS deorbit burn is targeted so as to place the earth's horizon at a specified point in the spacecraft window for monitoring purposes.
In summary, then, as regards launch aborts, the main things to remember are: (1) abort capability in one mode or another is available throughout the entire launch phase on a continu­ous basis; (2) contingency orbit insertion followed by de­orbit is always prime when it is available.
ABORTS FROM EARTH PARKING ORBIT COAST
This mission phase consists of coasting in a circular earth parking orbit from earth orbit insertion to the initiation of translunar injection, a duration which is usally one orbit or longer. As might be expected for this phase the abort procedures being planned are very similar to those used in Projects Mercury and Gemini. ​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.
​====The other abort mode possible for this phase is a DPS coplanar deorbit burn. This mode, as shown in figure 6, is available throughout the entire phase, but obviously requires transposition and docking in earth parking orbit.
Figure 8 shows the basic difference between this mode and the previously described one. Note that the deorbit maneuver is such that the spacecraft passes through apogee in order to provide enough time for the crew to transfer from the LM to the CSM prior to reentry. Also, the horizon cannot be easily monitored during the abort burn due to the spacecraft docked configuration. Thus, this attitude restrictionis deleted for this mode. As for the S-IVB mode, the use of this mode is very improbable if the Service Module RCS deorbit proves feasible due to the same type of recontact problems upon reentry as would be experienced with the S-IVB deorbit mode.
An alternate DPS abort mode not shown on the summary chart is currently being investigated. This new DPS mode would consist of using the DPS to lower perigee to very near the atmosphere so that the resulting spacecraft trajectory would then be within Service Module RCS capability to deorbit. In other words, the procedure would require maneuvers by both the DPS and the Service Module RCS. This procedure would eliminate recontact problems during reentry, since the LM would be jettisoned between the DPS burn and the Service Module RCS burn.
Summarizing for this phase, one can say that more than ade-quate abort capability exists from strictly a performance standpoint with essentially three independent propulsion systems capable of providing the abort $\Delta \mathrm {V}$ required, continu­ously through the mission phase. However, the use of the S-IVB and DPS, as described in this section, would be very undesirable due to the recontact problems during reentry. If use of the Service Module RCS system to deorbit proves feasible, then the abort modes using the S-IVB or the DPS can essentially be eliminated from consideration.
Since the SPS mode is very similar to that planned for Projects Mercury, Gemini, and the early Apollo orbital flights, the detailed procedures and computer programs are already available and checked out for the ground com­puters. The SPS deorbit modes could be executed using either the onboard G&N system, the SCS system, or a strictly manual-type abort using visual attitude reference. The DPS abort mode would normally be executed using the onboard G&N system, although it could also be executed using the backup AGS system.
This phase consists of the S—IVB burn which injects the spacecraft on a highly elliptical trajectory to rendezvous with the moon. The duration of this phase is approximately 3′0 seconds for a typical lunar mission. What is or interest here is to consider abort capability from orbits which would result from a premature or early burnout of the S-IVB stage
Figure 9 shows the type of preabort orbits which would result from an S-IVB underburn during this phase or the mission. Note that the family of orbits are elliptical, having very nearly coincident lines of apsides with ever increasing apogee altitude up to and beyond lunar distance. The perigee altitude, however, remains relatively fixed, very near that of the original circular earth parking orbit altitude. The periods or these orbits, as shown in figure 10, vary all the way from 1½ hours to approximately 400 hours as burn time increases. Actually, for free—return translunar profiles, the moon's gravitation perturbs the trajectory resulting from nominal burnout such that return to earth requires much less than 400 hours. Note also in figure 10 that the period remains relatively small (less than 10 hours) for more than three quarters of the way through the burn.
Table II shows the propulsion systems available which are capable of performing abort maneuvers during one portion or another of this phase as well as the subsequent phase, translunar Coasti Note that there is a large $\Delta \mathrm {V}$ capability with the service propulsion system even when the LM is attached. Also, there is moderate capability available with the LM propulsion systems, although use or them requires transposition and docking prior to abort. Note, however, that very little Capability is available with the Service Module RCS. Because of this, aborts using the Service Module RCS are marginal at best, as will be shown later.
Figure 11 presents a summary of the abort modes available for translunar injection. Note that redundant abort capability exists throughout the entire phase and even double redundancy for the latter part or the burn. The first and primary abort mode shown in the summary chart consists of a single burn to return the spacecraft directly to reentry, as shown sketched in Figure 12. This mode is available throughout the phase with either the SPS or DPS propulsion systems. Note also that the burn attitude is not constrained to be either coplanar or to enable horizon monitoring. A constrained attitude for aborts from the family of ellipses could result in excessive $\Delta \mathrm {V}$ penalties also prevent landing at a desired recovery area.
​Seven sub-modes exist which would be available from the ground for short in this one basic mode. In other words, there exists seven different classes or abort trajectories which will be possible with a single unconstrained atttude burn, There was really no requirement for this kind of flexibility in the previous phase—in otherword, during the earth parking orbit phase—because the spacecraft was always in the close proximity of the earth, and there was really no significantly different ways to return to reentry, Beginning, with this phase, however, the spacecraft could be in a preabort orbit which extends a consderable distance from the earth, This fact, combined with the large spacecraft performance capability, means that abort trajectories having significantly different characteristics as regards $\delta \mathrm {V}$, return time, and landing point are possible, Therefore, plans are being made to take advantage of this situation to provide greater flexibility in real time planning.
The Seven sub-modes available are as follows:
Figure 18 shows typical return times to the same primary recovery site as considered on the previous plot. As for the previous plot, the data is shown plotted as a function of trans lunar injection burn time and abort delay time. These data assume that the entire SPS capacity available for use. Note that although a few regions are available where return is in a few hours are possible, in general, returns are forced to at least one day in return time.
Summarizing the trans lunar injection phase, then, one can state that: (1) redundant abort capability exist continuously throughout the translunar injection phase; (2) spacecraft performance margin is great enough to allow considerable flexibility in selection of return times and choice of a landing area; (3) the only critical or marginal abort mode is when the Service Module RCS system must be relied upon to perform the abort maneuver. This mode would only be required in the event of an SPS failure as well as failure to perform transposition and docking to obtain the LM propulsion systems. Thus, the use of this mode would seem very improbable.
ABORTS DURING TRANSLUNAR COAST PHASE
This mission phase consists of the coast period from translunar injection burnout to the initiation of lunar orbit insertion, a period of approximately 62 to 74 hours for free-return trajectories. Since the nominal translunar coast trajectory is a free return to earth, the minimum ΔV to abort is essentially zero. Thus, as will be shown later, large ΔV capability margins exist which can be used in an abort situation to speed up the return to earth and/or control the point of landing.
Figure 19 presents a summary of the abort modes for the translunar coast phase. Note that redundant abort capability exists throughout the entire phase due to the presence of three independent propulsion systems.
​The first abort mode listed is a direct abort available with either the SPS or LM DPS propulsion systems. The basic features of this abort mode are shown sketched in figure 20. This mode consists of a single unconstrained attitude abort burn which returns the spacecraft directly to reentry without circumnavigating the moon. Shown sketched, are the two extreme types of returns possible in this mode--time critical and fuel critical. A fuel critical return usually passes through apogee following the abort maneuver; whereas, a time critical or a fast return usually does not pass through apogee following the abort maneuver. All of the seven sub-modes described previously for a translunar injection phase are available for this mode as well as for all of the other modes for this phase. If the SPS is available, this mode of abort produces the fastest possible returns to the earth for aborts performed prior to reaching approximately the lunar sphere of influence. This is the reason why this particular mode is considered prime for approximately the first three-fourths of the coast period from the earth to the moon.
The DPS direct abort mode, as shown in the summary chart, is available as a backup mode to the SPS direct mode for approximately the first two-thirds of the coast period from the earth to the moon. Use of this mode during the latter portion of the translunar coast phase would result in excessively long return times. Thus, the use of this mode as a backup during this latter period is not considered as a possibility. As will be shown in more detail later, the LM DPS direct abort mode, when used in the event of an SPS failure, will result in the fastest possible return time only if the abort maneuver is performed during approximately the first 25 hours of translunar coast.
The next abort mode listed in the summary chart consists of delaying the abort maneuver until the vicinity of pericynthion or perilune is reached, as shown sketched in figure 21. This mode of abort is available continuously throughout the entire translunar coast phase with either the SPS, DPS, or marginally with the Service Module RCS system. If the SPS is available, this mode of abort will generally result in the fastest possible return to the earth for abort decisions made near the sphere of influence or thereafter. For this reason, it is shown in the summary chart as the prime mode of abort for this portion of the translunar phase. In the event of an SPS failure, the LM DPS propulsion system used in this mode will result in the fastest possible return time for abort decisions made approximately 25 hours or later out along the translunar coast trajectory. In the event of an SPS failure and failure to obtain the LM DPS propulsion systems, the Service Module RCS would be marginally available for this mode of abort due to the fact that the nominal translunar coast is a free-return trajectory.
​The third abort mode listed in the summary chart is a circumlunar abort available with the SPS, DPS, or marginally with the Service Module RCS continuously throughout this mission phase. Figure 22 shows the basic features of this abort mode to be a single unconstrained burn performed at some point along the translunar coast trajectory which returns the spacecraft to the earth after circumnavicating the moon. The altitude of pericynthion or perilune is allowed a certain amount of freedom in order to obtain landing area control upon return to earth. This mode of abort is not as yet thoroughly understood. Preliminary analysis indicates that this type of abort many times produces the absolute minimum ΔV required to return to a particular landing area, especially if the translunar coast profile is a non-free-return type. However, for just as many cases the minimum ΔV required to return to a particular recovery area has been found to be minimized by use of delay to pericynthion abort mode. A complete understanding of this effect is currently under investigation.
Figure 23 shows a special case of the circumlunar abort mode. This particular case consists of merely using mid-course corrections to correct back to the nominal free-return trajectory, assuming this type of return is the nominal profile. Use of the circumlunar abort mode in this fashion is essentially equivalent to the fuel critical unspecified area sub-mode described previously.
The last mode listed on the summary chart is a two-burn mode for the special case of when the translunar coast trajectory is on an impact course with the moon. As shown, this mode of abort is available continuously throughout the entire phase with either the SPS, DPS, or marginally with the Service Module RCS. Figure 24 shows the basic features of this abort mode. The first maneuver, usually a small one, is used to raise perilune or pericynthion altitude to an acceptable value. The second burn is then performed in the vicinity of perilune to return the spacecraft to earth in one of the seven sub-modes described previously. The most brobable need for this abort mode would be in the event of a badly executed second midcourse correction near the sphere of influence due to a G&N or an SPS failure.
Figure 25 shows a very significant characteristic f trans lunar coast abort trajertories, expecially as regards returns to a primary recovery site. As shown, the transfe angle from abort to reentry is relatively insensitive·t,ce time of abort or to the type of abort return. This angle varies from approximately 170° to 180° throughout the entire translunar coast phase. This fact, combined with the fact that the reentry ranging capability is also relatively fixed, ​means that the inertial position of landing is approximately fixed for a given return plane. This means that the abort trajectory problem is essentially a timing or a rendezvous problem. In other words, the return must be timed such that, as the spacecraft reaches its relatively fixed inertial landing point, the desired landing area is just rotating underneath. If the fastest possible abort return just misses rendezvous with the desired landing area, then a delay in landing time of nearly 24 hours results in order to allow the desired landing area to make a complete revolution to again reach the desired inertial rendezvous point. Thus, more than one solution is possible, but the possible solutions occur in 24-hour increments as regards landing time.
Figure 26 shows that this 24-hour landing effect is also present when the time of abort is added as an additional degree of freedom to the problem. This effect occurs because the inertial direction of the abort point changes very little with abort time along the translunar coast trajectory. An abort from point 1, as shown in figure 26, to a particular landing site might require relatively low energy in order to achieve rendezvous with a desired landing site. An abort at a later time, as represented by point 2, would require a higher energy return trajectory if the spacecraft is to rendezvous with the same landing site at the same time as the abort trajectory 1. Similarly, an abort trajectory performed at still a later time, point 3, would require even more energy resulting in an even faster return trajectory in order to account for the time lost in delaying the abort maneuver. Obviously, a limit will be reached as regards delay time when either the A.V required for abort or the velocity at reentry violates its respective limit. When this limiting or critical delay time is reached, an abort would be forced to a landing 24 hours later, as represented by point h in figure 26.
An interesting feature of this 24-hour effect for aborts to a particular landing site is that returns will either all be in daylight or all in darkness. In other words, if an abort solution to a primary recovery site such as Hawaii lands in darkness, a solution cannot be found for any other abort delay time or any other AV required which will cause landing at Hawaii to be in daylight.
Figure 27 shows this 2h-hour effect in the form of actual abort trajectory performance data computed for a typical translunar coast trajectory. Shown plotted is the time
of landing measured from translunar injection as a function of the time of abort measured from translunar injection. Data is shown for three primary recovery sites--Indian Ocean, Hawaii, and Bermuda. The data shown for aborts ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/362 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/363 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/364 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/365 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/366 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/367 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/368 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/369 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/370 ​Page:LunarLandingMIssionSymposium1966 1978075303.pdf/371 ​Figure 52 presents actual abort performance data for aborts to primary recovery sites for a typical transearth coast trajectory shown plotted is the time at landing as measured from transearth injection as a function of the time or abort measured from transearth in action. Data is shown for three primary recovery sites—Indian Ocean, Hawaii, and Bermuda, As shown, the nominal transearth trajectory is targeted to return to Hawaii at the time of landing or 90 hours as measured from transearth injection. Aborts with a time of landing 24 hours earlier to Hawaii can be performed during the first in hours of transearth coast. Nate also that the raster returns are available to alternate landing sites during the first us hours of transearth coast, Also shown is the availability of returns to an alternate (Indian Ocean) site with a later time of landing up through approximately the first 55 hours of transearth coast, These data, then, show considerable flexibility or abort capability to speed up the return and/or to change the point of landing. It is also of interest that returns to primary recovery sites still display the 24 hour effect, as in the previous phases.
Questions and Answers
EARTH RETURN ABORT CAPABILITIES
Speaker: Ronald L . Berry
1.Dr. Rees - Why not use the APS for a backup propulsion system?
ANSWER - Since the ascent engine is not gimballed, the possible c.g. offset effects cannot be controlled.
2.What is the RCS engine burn time limitation?
ANSWER - Specifications limit is 1000 seconds.
3.Dr. Mueller - Have methods to restart the SPS engine in the event of an early shutdown during transearth injection burn been investigated?
ANSWER - If the problem is a guidance or control problem, the engine can be restarted and controlled manually the problem is with the SPS engine itself, nothing can be done. This is one of the accepted risks in the program.
4.Dr. Haeussermann - Have we looked into using a DPS-SPS combination?
ANSWER - No. The main reason for using the DPS as backup is because of an SPS failure.
5.Mr. Richter - Can't we use ADS some of the time - isn't it load dependent?
ANSWER - The control authority is marginal in most cases for the APD since the APS engine is not gimballed.
6.Mr. Richter - Can we use two SPS burns for fast earth return transfer?
ANSWER - This is being looked into. This may not be a desirable way to achieve a gain in return time, since it could be dangerous. If the SPS did not fire the second time, you might exceed acceptable entry velocity or conditions.
​7.Mr. Green - When the DPS is used for backup propulsion, do we jettison the SPS propellants?
ANSWER - There is no capability to jettison SPS propellants .
8.Dr. Mueller - Have the procedures, etc., have been worked out for all of the abort possibilities?
ANSWER - No.
9.Dr. Von Braun - Why the difference in the translunar and transearth transit times?
ANSWER - The trans lunar phase uses a free return trajectory which limits the transfer time to a narrow band. The transearth time is primarily limited by the energy available and the amount of consumables remaining.
LAUNCH ABORT MODES
Fig. 1
LES ABORT SEQUENCE
Fig. 2
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