By JOACHIM P. KUETTNER, Ph. D., Chief, Saturn-Apollo Systems Integration Office, NASA George C. Marshall Space Flight Center; and EMIL BERTRAM, Chief, Special Projects Office, NASA Launch Operations Center




[69] The Mercury-Redstone launch vehicle was used for the first United States ballistic manned space flights. As a prelude to the orbital flight program, the Mercury-Redstone missions provided an opportunity to evaluate the performance of the Mercury spacecraft, the reactions of the astronauts to brief periods of space flight, and the launch and recovery operations. The first steps toward man-rating a tactical missile were made in a series of design changes and modifications based on ground and flight testing. This paper describes development of the first U.S. manned launch vehicle, including the abort system, the reliability programs necessary for pilot safety, and the performance of the Mercury-Redstone space vehicle.




The Mercury-Redstone launch vehicle was the United States' first manned launch vehicle. However, it is only the first of a series of launch vehicles which will exhibit an increasing capability in manned space payloads.


By early 1959, several decisions were made in regard to the performance required of a launch vehicle needed for the first phase of the manned flight program. The vehicle had to have both the reliability and performance to place a manned, 2-ton payload safely into a suborbital trajectory in which at least 5 minutes of weightlessness would be experienced and an apogee of at least 100 nautical miles would be attained. In addition, the vehicle had to be available in time to support the desired flight schedule. These requirements narrowed the choice to launch vehicles which had already been developed for a military mission.


At this time, two surplus Jupiter C missiles were available from the Army Ballistic Missile Agency (ABMA). The Jupiter C was an advanced version of the Redstone, a tactical military missile with a record of over 50 successful flights to verify its reliability. The original Redstone could not meet the mission requirements; however, the Jupiter C had elongated propellant tanks, a lighter structure, and the required performance for Mercury. The Jupiter C launch vehicle had been used for conducting reentry studies and placing the first U.S. satellite, Explorer I, into orbit.


Therefore, the Redstone vehicle, in its Jupiter C modification, satisfied the basic Mercury suborbital requirements of availability and performance.


However, the Jupiter C did not incorporate all the necessary safety features; and further adaptation was necessary for use as a manned launch vehicle. This development, which is sometimes referred to as "man rating," had as its three major guidelines safety during launch, satisfactory operation from a human-factors standpoint, and adequate performance margins.


The actual adaptation took place in three phases: basic modifications, modifications after ground tests, and modifications after flight tests. Although there were specific hardware changes during the development, the basic man-rating program and design concepts did not require major alteration.


Basic Vehicle Modification


As noted, some basic modification was necessary to adapt the Jupiter C to the Mercury mission requirements. The required modifications [70] and additions made the new Mercury-Redstone launch vehicle physically distinguishable from both the Redstone and Jupiter C missiles. Figure 4-1 illustrates the differences between these configurations. It should be noted that each successive version of the original Redstone was progressively longer.


Figure 4-1. Comparison of the three Redstone missiles.


To meet performance requirements, use of the elongated Jupiter G tanks was necessary. These tanks give the Mercury-Redstone launch vehicle a nominal engine burning time of 143.5 seconds, 20 seconds more than the original Redstone vehicle. This greater burning time required the addition of a seventh high-pressure nitrogen tank to pressurize the larger fuel tank and an auxiliary hydrogen peroxide (H202) tank to power the engine turbopump.


To decrease the complexity for the basic Mercury-Redstone, three changes were made:

(1) The Redstone stabilized platform (ST80) were replaced by the LEV-3 autopilot for vehicle guidance. The LEV-3 system, although less complex, was more reliable and met the guidance requirements of the Mercury Redstone mission

(2) The aft unit, containing the pressurized instrument compartment, and adapter were permanently attached to the center tank assembly. In the tactical version, these units separated with the payload to provide terminal guidance.

(3) A short spacecraft adapter, including the spacecraft-launch-vehicle separation plane, was supplied by the spacecraft contractor. This arrangement simplified the interface coordination.


To prevent major changes midway in the program, the engine was immediately changed from the A-6 to the A-7 model. The A-6 engine was scheduled to be phased out, and a shortage of hardware was expected to occur during the Mercury-Redstone program. This early changeover avoided a foreseeable problem area but required an accelerated test program.


For the Mercury-Redstone launch vehicle, alcohol was chosen as the fuel. Although Jupiter C had used unsymmetrical diethyltriamine (UDETA) for greater performance, it toxicity was higher than that of alcohol an was considered to be undesirable for manned flights. However, the selection of alcohol led to a problem with the important jet controls vanes because of the extended burning tire; which caused greater erosion of these vanes"; Hence, a program was initiated to select jet vanes of the highest quality for use in Mercury.


The prevalves were deleted from the Mercury-Redstone launch vehicle in order to increase mission success. These valves had been! used in the tactical missiles between the propellant tanks and the main propellant valves to; prevent possible fuel spillage in the event of main valve failure. However, failure of the prevalves to remain open in flight would have resulted in a mission abort.


To provide for maximum crew safety, an automatic inflight abort- sensing-system was added to the launch vehicle and an emergency egress operation was established for the launch complex. These factors were primary considerations in man-rating the Redstone and are discussed in greater detail later.


The Mercury-Redstone was aerodynamically less stable than the standard Redstone. Because of the unique payload characteristics and [71] the elongated tanks, the Mercury-Redstone was expected to become unstable in the supersonic region approximately 88 seconds after lift-off. (See fig. 4-2.) To compensate for this instability to some degree, 687 pounds of ballast were added forward of the instrument compartment.


Changes were also necessary because of the decreased lateral bending frequencies. The configuration and payload changes reduced the Mercury- Redstone bending frequencies to one fourth those experienced by the standard Redstone. (See fig. 4-3.) As a result, resonance problems appeared during both ground and flight testing. The second bending mode had to be filtered out of the control system to prevent feedback.


Figure 4-2. Center-of-gravity and center-of-pressure location of Mercury-Redstone during time of flight.


In all, a total of 800 changes were made before the Mercury-Redstone project was completed. The major modifications just described, as well as many minor changes beyond the scope of this paper, resulted in a reliable man-rated vehicle.


Figure 4-3. Mercury-Redstone lateral bending modes.


Abort System Description


Even though the vehicle was expected to perform properly, a launch- escape system was required for maximum crew safety as long as a catastrophic launch-vehicle failure remained a possibility. Therefore, an automatic inflight system was developed which supplied an abort signal to the spacecraft in the event of an impending catastrophic failure of the launch vehicle. This signal caused engine cut-off, escape-rocket ignition, and spacecraft separation. This cut-off mode was in addition to those sent when the mission conditions were achieved and in the event an emergency command destruct signal had to be sent. Because the vehicle was to be manned, the destruct signal had a built in 3-second delay to allow time for adequate spacecraft separation. The abort system, shown in figure 4-4, sensed and was activated by: unacceptable deviations in the programed attitude of the launch vehicle, excessive turning rates, loss of thrust or loss of electrical power.


72] Figure 4-4. Block diagram of Mercury-Redstone automatic abort sensing system.


The criteria for the abort system were based on an evaluation of over 60 Redstone and Jupiter C flights and a failure-mode analysis. The number of parameters was kept at a minimum, since an overly complicated system could result in little improvement, if any, in overall flight safety. A selection of those parameters which would reflect the operation of only vital systems was therefore required. Hence the abort system sensed primarily output or downstream parameters, each of which were then representative of many different types of failures.


For example, a sudden change in the attitude of the vehicle indicated trouble in the control system, regardless of the source of this trouble. It could be the result of a failure in the control computer or some mechanical system or the limits of controllability having been exceeded. By establishing critical values for pitch, yaw, and roll angle, a variety- of problems, including the unstable "flip-over'' with a subsequent explosion, could be predicted in time for a safe abort. As other examples, loss of thrust, rough combustion, and an impending explosion could be sensed from variations in the combustion chamber pressure. Finally, a loss in electrical power or of the electrical interface between the spacecraft and launch vehicle could be effectively sensed.


As shown in figure 4-4, the abort system circuitry was designed to include adequate redundancy. The combustion chamber pressure (Pc) switches were wired in parallel to assure an abort capability even if one sensor failed. Since the predominant failure mode of electrical voltage sensors is opposite that for a pressure switch, the relays controlled by the voltage sensors were connected in series. Although single sensor monitored pitch, yaw, and roll attitudes, as well as pitch and yaw attitude rates, redundancy was implicit for these attitude and rate measurements because of their interdependency.


To supply the necessary timing functions the abort system, relay interlocks were used to prevent arming of the abort system prior to lift- off and to disarm the system at normal shutdown. The Pc switches were armed after engine start and disabled prior to normal shutdown. Here, additional relays also provided circuit redundancy and lock-in of the abort signal.


Time is a critical factor in the abort procedures, and the method of abort initiation is completely dependent on it. Because some launch vehicle failures could very rapidly result in a catastrophe, the abort was designed to be automatically initiated. Since some failure would not cause an immediate catastrophe, manual backup was incorporated. The astronaut, blockhouse, mission control center, and range safety could initiate an abort during specifically assigned flight periods, some of which overlapped.


Nominal Mission Profiles


The Mercury-Redstone launch vehicle whose nominal mission profile is shown in figure 4-5, accelerated the Mercury spacecraft into a suborbital flight at a nominal speed of approximately 6,460 feet per second. At launch-vehicle-spacecraft separation the flight-path angle was 41.80°, the altitude, 200,000 feet, and the Mach number, 6.30. The maximum acceleration at cut-off was 6.3g.


In figure 4-5, several important launch vehicle sequencing points are indicated. A circuit permitting automatic engine cut-off prior to abort was activated 30 seconds after lift-off. Prior to this time, this circuit was disabled be cause cut-off in the first 30 seconds would have resulted in an impact of the launch vehicle on land which was undesirable; therefore, only the range safety officer could initiate an engine shutdown. To prevent an early jettisoning of the [73] escape tower the normal shutdown circuitry was not armed until 129.5 seconds. At 131 seconds the velocity cut-off accelerometer was armed. This arming occurred 12 seconds before nominal expected engine cut-off time to allow for higher-than-expected launch-vehicle performance for n non-optimum mixture ratio, which could result in premature propellant depletion. The chamber pressure sensors to the automatic abort system were deactivated at 135 seconds, thus preventing an abort signal at the time of cut-off. Both cut-off activation and pressure switch deactivation were originally scheduled to occur at 137.5 seconds, but as a result of the early shutdown of MR-2, the times indicated in the figure were selected for all subsequent flights.


At engine shutdown, nominally at 143 seconds, the abort system was deactivated and the escape tower jettisoned. Spacecraft separation occurred 9.5 seconds after shutdown to allow for thrust tail-off.


Figure 4-5. Mercury-Redstone powered flight sequence.


Reliability, Testing, and Quality Assurance


As mentioned earlier, the basic launch vehicle had a history of 69 flights prior to the first manned flight upon which to base failure-mode and reliability prediction. Two such predictions were made. The first prediction used the record of all Redstone, Jupiter C, and Mercury Redstone development and qualification flights. The second prediction used an artificial Redstone configuration composed of individual components flow n at different times on previous flights.


To find the weak spots in the total vehicle, large subsystems were submitted to a special reliability test program. All major missile sections and the systems contained in each were vibrated under temperature and humidity conditions simulating the actual environments of transportation, prelaunch, and flight. Bending and compression loads were applied up to 150 [74] percent of maximum flight loads, thereby establishing positive margins of safety. When trouble spots were found, individual component testing was followed up with additional systems tests.


Figure 4-6 shows the vehicle contractor's combined environmental test facility. This facility applied flight vibrations and rigid body motions up to 4g at 2,000 cps simultaneously with temperatures up to 115° F. This testing proved the importance of investigating the interaction of all component masses.


combined vehicle test apparatus

Figure 4-6. The contractor's combined vehicle motion and vibration test stand.


In addition, structural flight simulation, spacecraft-launch-vehicle interface compatibility, clamp-ring operation, and static firing tests were made. Figure 4-7 shows the mating of the spacecraft and launch vehicle prior to a noise and vibration test conducted at the NASA Marshall Space Flight Center (MSFC).


Quality assurance procedures were relatively mole refined than for the tactical vehicle because of the stress placed on crew safety. An awareness program required that every Mercury system assembly carry a special Mercury stamp indicating that it had passed special inspections and that all personnel involved in its manufacture and assembly were aware of the quality expected. Particular attention was paid the areas involving soldering techniques, welding repairs, and preparation of instructions.


testing simulator

Figure 4-7. Static firing, noise and vibration stand.


Changes Resulting From Ground Tests


During the vibration test program, sever components failed or were damaged. The components included an engine piping elbow, an H2O2 bottle bracket, the abort-rate switch mounting bracket, wires in the roll-rate switch and an antenna mounting stud. Similar problems occurred in other components. The success of the modifications proved the value o total system testing.


Since the A-7 engine was new, extensive firings were made. During these firings, a instability was discovered at 500 cps and eliminated through a modification to the fuel injector. Investigation as to the source of another low-frequency oscillation eventually led to the discovery that the static test tower was at fault. Modification of the static test tower subsequently removed it as a trouble source.


Checkout and Launch Operations


Prior to shipment to the launch site at Cap Canaveral, the Mercury- Redstone abort system was checked by introducing simulated malfunctions and evaluating the abort system response.


[75] The first three launch vehicles were also carefully tested for compatibility with the spacecraft at MSFC.


At Cape Canaveral, the Mercury-Redstone countdown was conducted in two parts with a rest period in between to reduce fatigue of the launch crew. Lox loading was scheduled for completion at 180 minutes prior to lift-off to minimize the possibility of an additional 12-hour delay for lox tank purging and drying during the recycle time in the event of a launch cancellation after lox loading. The astronaut was to be inserted into the spacecraft after lox loading at approximately 120 minutes prior to lift off. A period of 4 hours was considered to be a tolerable time between astronaut insertion and lift-off to accommodate possible holds in the countdown.


Emergency Egress and Pad Abort


Special astronaut safety precautions were required after insertion since the launch vehicle was already fueled; therefore, launch pad emergency egress procedures were developed. A study (see fig. 4-8) to determine the best mode to retrieve an incapacitated astronaut indicated the blockhouse- controlled service structure would provide the most expeditious escape. If, however, he were able to exit without help,


Figure 4-8. Time study of astronaut emergency egress.


[76] he could use the pad escape tower, or "Cherry Picker," shown in figure 4-9. The cab of this specialized escape equipment, which was permanent but extendable, was stationed near the spacecraft hatch until just prior to lift-off. Utilization of this escape device was combined with the use of fire trucks, an armored personnel carrier (M-113), and rescue teams for exit from the pad area. In case of a pad abort, recovery procedures and vehicles, including army helicopters and amphibious craft, were organized and prepared to assist.


crane type apparatus to service spacecraft

Figure 4-9. MR-3 with "Cherry Picker" and remote controlled service structure.


System Modifications Resulting From Flight Operations


Problem areas revealed during the qualification flight-test program (MR- 1, MR-2, MR-BD) lead to the following modifications:

(1) The MR-1 launch attempt proved the need for ground-negative until all other electrical connections were separated. Thus a ground strap was added. This strap is shown in figure 4-10.

(2) A scale-factor error resulting from an excessive pivot torque on the LEV-3 longitudinal integrating accelerometer caused the MR-1A launch vehicle to experience a cut-off velocity exceeding the nominal value by about 260 feet per second. Use of softer wire and the relocation of the electrical leads eliminated the problem.


Figure 4-10. Mercury-Redstone ground strap.


(3) As a backup to the integrating accelerometer fix, a time-based cut- off signal was established at 143 sec for the MR-2 and MR-BD (booster development) missions. These later flights proved that the accelerometer functioned properly, and use of the cut-off timer was discontinued.

(4) The thrust controller on MR-2 failed wide open causing lox depletion 0.5 second before; deactivation of the abort Pc switches and before integrating accelerometer arming, which could have prevented this trouble. To prevent a simiIar occurrence on the remaining flights, velocity cut-off arming and Pc abort switch disarm were separated in time. Velocity cut-off arming was advanced to 131 sec to take care of earlier than predicted cut-off velocity, while Pc disarming was set at 135 sec, keeping the combustion chamber pressure abort capability as long as possible, but removing this capability early enough to take care of a high propellant consumption rate.

(5) Flights MR-1A, MR-2, and MR-BD experienced momentary roll rates approximately twice that of the earlier Redstone vehicle (~ 8°/sec as against ~ 4°/sec-abort limits were 12°/sec). Since the missile was not [77] subject to damage at this rate, the roll-rate abort sensor was deleted after MR-BD to increase mission success. The roll attitude angle abort limit of 10° was retained

(6) An interaction of the second bending mode with the yaw and pitch axis control required the addition of a network filter to reduce control loop gain between 6 and 10 cps. The interaction was noted on flights MR- 1A and MR-2 and is illustrated in figure 4-11.


Figure 4-11. Second bending mode oscillations in yaw toward end of MR- 1A flight.


(7) During MR-1A, MR-2, and MR-BD, undesirable vibrations in the adapter and instrument compartment were evident. On MR-3 these were dampened with 340 pounds of lead-impregnated plastic compound added to the bulkhead and walls of the section. The weight of this compound was substituted for an equal amount of ballast weight. Fourteen longitudinal stiffeners were also added to the internal skin surface. These improvements are depicted in figure 4-12. Since Astronaut Shepard still noted considerable vibrations during powered flight in MR-3, an additional 102 pounds of the dampening compound, X306, were added to the instrument compartment of MR-4. The summation of these changes resulted in the Mercury-Redstone shown in figure 4-13.


Flight Results


Three qualification flights were conducted for the Mercury Redstone flight series. MR-1 was launched on November 21, 1960. After rising a few inches, it settled vertically back on the launcher. It proved the need for careful examination of electrical circuitry and led to the addition of a strap for proper electrical grounding.


The sequence of events which led to MR-1's difficulties started during the lift-off when the power and control connectors did not disconnect simultaneously. Because of mechanical adjustments the power plug disconnected 29 milliseconds prior to the control plug. This permitted part of a 3-amp current, which would have normally returned to ground through the power plug, to pass through the "normal cutoff" relay and its ground diode. The cut-off terminated thrust and jettisoned the escape tower.


The spacecraft did not separate from the launch vehicle because the g- load sensing requirements in the spacecraft were not met. "Normal cut- off'' started a 10-second timer which, upon its expiration, was supposed to signal separation if the spacecraft acceleration was less than 0.25g. (This sequencing was designed to minimize the occurrence of a spacecraft launch-vehicle recontact. However, MR-1 had settled on the pad before the timer expired and the g-switch sensing 1g blocked the separation signal.)


The barostats properly sensed that the altitude was less than 10,000 ft and therefore actuated the drogue, main, and reserve parachutes in the proper sequence. The reserve parachute was released because no load was sensed on the main parachute load sensors. To prevent this failure from recurring, engine pressure was monitored and, if normal at 129.5 seconds, the normal booster cut-off signal path to the spacecraft was armed.


Following the MR-1 attempt, the spacecraft was refurbished and mated to a new launch vehicle, scheduled to be launched as MR-1A. The MR-1A space vehicle successfully accomplished the MR-1 mission objectives on December 19. 1960. The launch was slightly compromised by a scale- factor error in the longitudinal integrating accelerometer which caused cut-off velocity to be 260 feet per second higher than normal This higher velocity caused the spacecraft to experience somewhat higher reentry deceleration . During the flight, all measured abort parameters remained below the limits and the abort system functioned as expected.


MR-2, launched January 31, 1961, carried chimpanzee named "Ham." On this flight, the


78] Figure 4-12. Installation of dampening compound in instrument compartment and adapter section for Mercury-Redstone 4 (MR-4).


thrust controller ran above nominal resulting in propellant depletion 0.5 second before abort pressure sensor deactivation. The abort system was able to sense this early shutdown and aborted the spacecraft. The above normal cut-off velocity, combined with the thrust of the escape motor caused the spacecraft to land well beyond the intended recovery area. The simple timing changes explained previously were made to take care of higher propulsion system tolerances.


MR-BD was launched on March 24,1961, to evaluate a filter network added in the launch vehicle control circuit and modifications incorporated to eliminate the overspeed condition experienced on MR-1A and MR-2. The filter network was intended to dampen the effect of the second bending mode frequency (6 to 10 cps) on the pitch and yaw loop. The flight went exactly as expected and proved the effectiveness of this change.


MR-3 was the first manned flight. With Astronaut Alan Shepard as the pilot, the spacecraft lifted off at 9:34 a.m. e.s.t. on May 5, 1961. All objectives assigned to the launch vehicle were successfully accomplished and no system malfunction occurred. During powered flight,


79] Figure 4-13. Mercury-Redstone configuration.


the astronaut reported buffeting. However, telemetry data indicated lower vibrations than on earlier flights. To reduce these vibrations, additional dampening material was added to the instrument compartment prior to the remaining flight.


Concluding the Mercury-Redstone program was MR-4 carrying Astronaut Virgil I. Grissom in the second manned suborbital space flight. Again, all launch-vehicle systems worked properly and all objectives were achieved. Improved vibration reports indicated that the additional dampening material added to the instrument compartment proved effective.


The Mercury-Redstone flight program was concluded on a positive note with the successful MR-4 mission on July 21, 1961. The first manned flight into space had been accomplished by MR-3 in just over 2 1/2 years from the project's initiation. The initial objectives of providing space flight familiarization and training for astronauts had been accomplished. The spacecraft was exposed briefly to space flight conditions. Of equal importance was the invaluable training of the ground crew in the preparation, launching, and the recovery of a manned spacecraft.


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