The Third Mission:

Testing the LM in Earth Orbit

3 March–13 March 1969





Apollo 9 was a Type D mission, a lunar module manned flight demonstration in Earth orbit. It was the first manned test of the “lunar ferry” that would put astronauts on the Moon. A lunar module had first flown without a crew during Apollo 5 on 22 January 1968.


Many of the LM tests on Apollo 9 would exceed conditions expected in a lunar landing. To ensure that major objectives would be accomplished if Apollo 9 ended early, the schedule for the first half of the mission also included more work for the crew than the schedule of either Apollo 7 or Apollo 8.


The primary objectives were:


  • to demonstrate crew, space vehicle, and mission support facilities performance during a manned Saturn V mission with command and service modules and lunar module;


  • to demonstrate lunar module crew performance;


  • to demonstrate performance of nominal and selected backup lunar orbit rendezvous mission activities; and


  • to assess command and service module and lunar module consumables.


To meet these objectives, the lunar module was evaluated during three separate piloting periods that required multiple activation and deactivation of systems, a situation unique to this mission.


The crew members were Colonel James Alton McDivitt (USAF), commander; Colonel David Randolph Scott (USAF), command module pilot; and Russell Louis “Rusty” Schweickart, lunar module pilot.


Selected in the astronaut group of 1962, McDivitt had been command pilot of Gemini 4. Born 10 June 1929 in Chicago, Illinois, he was 39 years old at the time of the Apollo 9 mission. McDivitt received a B.S. in aeronautical engineering from the University of Michigan in 1959. His backup for the mission was Commander Charles “Pete” Conrad, Jr. (USN).


Scott had been pilot of Gemini 8. Born 6 June 1932 in San Antonio, Texas, he was 36 years old at the time of the Apollo 9 mission. Scott received a B.S. from the U.S. Military Academy in 1954 and an M.S. in aeronautics and astronautics from the Massachusetts Institute of Technology in 1962. He was selected as an astronaut in 1963. His backup was Commander Richard Francis “Dick” Gordon, Jr. (USN).


Schweickart, a civilian, was making his first spaceflight. Born 25 October 1935 in Neptune, New Jersey, he was 33 years old at the time of the Apollo 9 mission. Schweickart received a B.S. in aeronautical engineering in 1956 and an M.S. in aeronautics and astronautics in 1963 from the Massachusetts Institute of Technology. His backup was Commander Alan LaVern Bean (USN).


The capsule communicators (CAPCOMs) for the mission were Major Stuart Allen Roosa (USAF), Lt. Commander Ronald Ellwin Evans (USN), Major Alfred Merrill Worden (USAF), Conrad, Gordon, and Bean. The support crew were Major Jack Robert Lousma (USMC), Lt. Commander Edgar Dean Mitchell (USN/Sc.D.), and Worden. The flight directors were Eugene F. Kranz (first shift), Gerald D. Griffin (second shift), and M.P. “Pete” Frank (third shift).


The Apollo 9 launch vehicle was a Saturn V, designated SA-504. The mission also carried the designation Eastern Test Range #9025. The CSM was designated CSM-104 and had the call-sign “Gumdrop,” derived from the appearance of the command module when it was transported on Earth. During shipment, it was covered in blue wrappings that gave it the appearance of a wrapped gumdrop. The lunar module was designated LM-3 and had the call-sign “Spider,” derived from its arachnid-like configuration.



Launch Preparations


The launch was originally scheduled for 28 February 1969, and the terminal countdown had begun for that launch at 03:00:00 GMT on 27 February at T-28 hours. However, one-half hour into the scheduled 3-hour hold at T-16 hours, the countdown was recycled to T-42 hours to allow the crew to recover from a mild viral respiratory illness. The count was picked up at 07:30:00 GMT on 1 March.


A low-pressure disturbance southwest of Cape Kennedy in the Gulf of Mexico was the principal cause of overcast conditions. At launch time, stratocumulus clouds covered 70 percent of the sky (base 3,500 feet) and altostratus clouds covered 100 percent (base 9,000 feet); the temperature was 67.3° F; the relative humidity was 61 percent; and the barometric pressure was 14.642 lb/in2. The winds, as measured by the anemometer on the light pole 60.0 feet above ground at the launch site, measured 13.4 knots at 160° from true north.



Ascent Phase


Apollo 9 was launched from Kennedy Space Center Launch Complex 39, Pad A, at a Range Zero time of 16:00:00 GMT (11:00:00 a.m. EST) on 3 March 1969. The planned launch window for Apollo 9 extended to 19:15:00 GMT.


Between 000:00:13.3 and 000:00:33.0, the vehicle rolled from a launch pad azimuth of 90° to a flight azimuth of 72°. The S-IC engine shut down at 000:02:42.76, followed by S-IC/S-II separation and S-II engine ignition. The S-II engine shut down at 000:08:56.22, followed by separation from the S-IVB, which ignited at 000:09:00.82. The first S-IVB engine cutoff occurred at 000:11:04.66, with deviations from the planned trajectory of +2.86 ft/sec in velocity and -0.17 n mi in altitude.


The S-IC stage impacted at 000:08:56.436 in the Atlantic Ocean at latitude 30.183° north and longitude 74.238° west, 346.64 n mi from the launch site. The S-II stage impacted at 000:20:25.346 in the Atlantic Ocean at latitude 31.462° north and longitude 34.041° west, 2,413.2 n mi from the launch site.


The maximum wind conditions encountered during ascent were 148.1 knots at 264° from true north at 38,480 feet, with a maximum wind shear of 0.0254 sec-1 at 48,160 feet.


Parking orbit conditions at insertion, 000:11:14.66 (S-IVB cutoff plus 10 seconds to account for engine tailoff and other transient effects), showed an apogee and perigee of 100.74 by 99.68 n mi, an inclination of 32.552°, a period of 88.20 minutes, and a velocity of 25,569.78 ft/sec. The apogee and perigee were based upon a spherical Earth with a radius of 3,443.934 n mi.


The international designation for the CSM upon achieving orbit was 1969-018A; the S-IVB was designated 1969-018B. After undocking, the LM ascent stage would be designated 1969-018C and the descent stage 1969-018D.



Earth Orbit Phase


After post-insertion checkout, the CSM was separated from the S-IVB stage at 002:41:16.0. The adapter panels that housed the LM and shielded it from the rigors of launch was then jettisoned. The CM was turned so its apex, holding the docking probe, faced the LM. Docking with the LM was completed at 003:01:59.3.


After docking, the commander and lunar module pilot began preparations for their eventual entry into the LM. They pressurized the tunnel between the two spacecraft, and with the aid of the CMP, removed the CM hatch and checked the latches on the docking ring to verify the seal. Then they connected the electrical umbilical lines that would provide power to the LM while docked to the CM. The hatch was then replaced.


At 004:08:09, an ejection mechanism, used for the first time, ejected the docked spacecraft from the S-IVB.


Following a separation maneuver, the S-IVB was restarted at 004:45:55.54 and burned for 62.06 seconds. Ten seconds later, the S-IVB entered a 1,671.58 by 105.75 n mi intermediate coasting orbit that would allow the engine to cool down sufficiently prior to a restart within one revolution. The period of the orbit was 119.22 minutes, the inclination was 32.302°, and the velocity at insertion was 27,753.61 ft/sec.


At 005:59:01.07, the crew performed the first of eight service propulsion firings, a 5.23-second maneuver that raised the CSM/LM orbit to 127.6 by 111.3 n mi.


The third and final S-IVB ignition at 006:07:19.26 was a 242.06-second maneuver to demonstrate restart capability after the 80-minute coast and to test the engine performance under “out-of-specification” conditions. It also provided better ground tracking lighting conditions for the upcoming rendezvous. The escape orbit was achieved 10 seconds after S-IVB engine cutoff, and the velocity was 31,619.85 ft/sec. S-IVB performance was not as predicted due to various anomalies, including the failure of an LH2 and LOX dump. The LH2 dump through the engine could not be accomplished due to loss of pneumatic control of the engine valves. The LOX dump was not performed due loss of engine pneumatic control during the third burn. The LOX tank was satisfactorily safed by utilizing the LOX non-propulsive venting system.


The third ignition also served to place the S-IVB into a solar orbit with an aphelion and perihelion of 80,280,052 by 69,417,732 n mi, an inclination of 24.390°, an eccentricity of 0.07256, and a period of 325.8 days.


Crew activity on the second day was devoted to systems checks, pitch, roll and yaw maneuvers, and the second, third, and fourth service propulsion system burns while docked to the LM. The second burn, a 110.29-second maneuver at 022:12:04.07, raised the orbit to 192.5 by 110.7 n mi. The third burn, at 025:17:39.27, lasted 279.88 seconds. It raised the orbit to 274.9 by 112.6 n mi and lightened the spacecraft so that it could be controlled by the reaction control system engines later in the mission and be in a better rescue position for rendezvous activities. During these two burns, tests were made to measure the oscillatory response of a docked spacecraft to provide data to improve the autopilot response for this configuration. The fourth burn, at 028:24:41.37, was a 27.87-second phasing maneuver to shift the node east and put the spacecraft in a better position later for lighting, braking, and docking.


On the third day, at 043:15, the lunar module pilot transferred to the LM to activate and check out the systems. The commander followed at 044:05. The LM landing gear was deployed at 045:00.


At about 045:52, the commander reported that the lunar module pilot had been sick on two occasions and that the crew was behind in the timeline. For these reasons, the extravehicular activity was restricted to one daylight pass and would include only the opening of the hatches of the CM and LM. It was also decided to keep the lunar module pilot connected to the environmental control system hoses.


After communication checks for both vehicles, a five-minute television transmission was broadcast at 046:28 from inside the LM. The camera was trained on the instrument displays, other features of the LM interior, and the crew. The picture was good, but the sound was unsatisfactory.


The LM descent engine was fired for 371.51 seconds at 49:41:34.46 with the vehicles still docked. Attitude control with the digital autopilot and manual throttling of the descent engine to full thrust were also demonstrated. Crew transfer back to the CM began at 050:15, and the LM was deactivated at 051:00. The fifth service propulsion system firing, 43.26 seconds in duration, occurred at 054:26:12.27 to circularize the orbit for the LM active rendezvous. The resulting orbit was 131.0 by 125.9 n mi, compared to a desired circular orbit of 130.0 n mi, but it was considered acceptable for the rendezvous sequence.


Extravehicular operations were demonstrated on the fourth day of the mission. The plan was for the lunar module pilot to exit the LM, transfer to the open hatch of the CM, and then return. This plan was abbreviated from 2 hours 15 minutes to 39 minutes because of several bouts of nausea experienced by the lunar module pilot on the preceding day and because of the many activities required for rendezvous preparation.


The LM was depressurized at 072:45 and the forward hatch opened at 072:46. The lunar module pilot began his egress to the forward platform at 72:59:02, feet first and face up, and completed egress at 073:07. He was wearing the extravehicular mobility unit backpack, which provided communications and oxygen; it also circulated water through the suit to keep him cool. His only connection to the LM was a 25-foot nylon rope to keep him from drifting into space. He secured his feet in the “golden slippers,” the gold-painted restraints affixed to the surface outside the hatch, called the “front porch” by the astronauts, where he remained while outside the LM.


During this same period, the command module pilot, dependent on CSM systems for life support, depressurized the CM and opened the side hatch at 073:02:00. He partially exited the hatch for observation, photography, and retrieval of thermal samples from the side of the CM. The samples were missing, so he retrieved the service module thermal samples at 073:26. The lunar module pilot retrieved the LM thermal samples at 073:39. Three minutes later, he began an abbreviated evaluation of translation and body-attitude-control capability using the extravehicular transfer handrails. The initially planned hand-over-hand trip from the LM to the CM was not made. During this period, the lunar module pilot also completed 16 mm and 70 mm photography of the command module pilot’s activities and the exterior of both spacecraft.


The lunar module pilot began his ingress at 073:45 and completed it at 073:46:03. The LM forward hatch was then closed, locked, and the LM was repressurized by 073:53. The CMP had already closed and locked the CM hatch, completing this task by 073:49:23. The CM was repressurized by 074:02:00.


The first images of the television transmission from inside the LM were received at 074:58:03 and the transmission was concluded with loss of signal at 075:13:13. Voice and pictures were both good, an improvement over the previous day’s transmission. The commander returned to the CM at 075:15, followed by the lunar module pilot at 076:55.


On the fifth day, the lunar module pilot transferred to the LM at 088:05, followed by the commander at 088:55, to prepare for the first LM free flight and active rendezvous.


The CSM was maneuvered to the inertial undocking attitude at 092:22. Undocking was attempted at 092:38 but the capture latches did not release immediately. Undocking occurred at 092:39:36, and the LM was rolled on its axis so that the CMP could make a visual inspection. A small separation maneuver at 093:02:54, using the service module reaction control system, placed the LM 2.0 n mi behind the CSM 45 minutes later. The maximum range between the LM and CSM was 98 n mi, achieved about halfway between the coelliptical sequence initiation and constant differential height maneuver.


During this maneuver, the LM engine ran smoothly until throttled to 20 percent, at which time it chugged noisily. The commander stopped throttling and waited. Within seconds, the chugging stopped. He accelerated to 40 percent before shutting down and had no more problems. The LM crew then checked their systems and fired the descent engine again to 10 percent. It ran evenly.


The first LM rendezvous phasing maneuver was executed at 093:47:35.4 with the descent propulsion system under abort guidance control. This maneuver placed the LM in a near equiperiod orbit with apogee and perigee altitudes 12.2 n mi above and below the CSM. The second maneuver was not applied; it was a computation to be used only in case of a contingency requiring a LM abort. The solution time was 094:57:53. The third rendezvous maneuver was executed at 095:39:08.06 and resulted in an LM orbit of 138.9 by 133.9 n mi.


Coelliptic sequence initiation was performed at 096:16:06.54, and the descent stage was jettisoned immediately after the start of reaction control system thrusting. The maneuver left the LM 10 n mi below and 82 n mi behind the CSM. The descent stage remained in Earth orbit until 03:45 GMT on 23 March, when it impacted the Indian Ocean off the coast of eastern Africa.


The resulting ascent stage orbit was 116 by 111 n mi. After coelliptic sequence initiation using the CSM reaction control system, rendezvous radar tracking was reestablished, but the CM was unable to acquire the ascent stage tracking light, which had failed. The constant differential height maneuver was performed at 096:58:15.0, using the ascent stage engine for the first time. The onboard solution for terminal phase initiation was executed at 097:57:59, creating an ascent stage orbit of about 126 by 113 n mi. Two small midcourse corrections were performed at 10 and 22 minutes after terminal phase initiation. Terminal phase braking began at 098:30:03, followed by stationkeeping, formation flying, photography, and docking at 099:02:26. The ascent stage had been separated from the CSM for 6 hours 22 minutes 50 seconds.


After docking, the crew transferred back to the CSM by 101:00. The ascent stage was jettisoned at 101:22:45.0, and the ascent engine fired for 362.3 seconds at 101:53:15.4 until oxidizer depletion. The final orbit for the ascent stage was 3,760.9 by 126.6 n mi, with an expected orbital lifetime of five years; however, entry occurred on 23 October 1981.


The sixth service propulsion burn, a 1.43-second maneuver at 123:25:06.97, had been postponed for one revolution because the reaction control translation required prior to ignition for propellant settling was improperly programmed. The maneuver, originally scheduled for 121:48:00, was an orbit-shaping retrograde maneuver to lower the perigee so that the reaction control system deorbit capability would be enhanced in the event of a contingency.


During the final four days in orbit, the crew conducted Earth resources and multispectral terrain photography experiments over the southern United States, Mexico, Brazil, and Africa. One objective, designated experiment S065, was to determine the extent to which multi-band photography in the visible and near-infrared regions from orbit could be effectively applied to the Earth resources disciplines.


The other objective was to obtain simultaneous photographs with four different film/filter combinations from orbit to assist in defining future multispectral photographic systems. The results were excellent. The quality and subject material exceeded that of any previous orbital mission and would aid in future program planning. The reasons for the excellent results were the amount of time available (four days so the crew could wait for cloud cover to pass); the orbital inclination of 33.6°, which permitted vertical and near-vertical coverage of areas never photographed before; sufficient reaction control propellants which allowed the crew to orient the spacecraft whenever necessary; the lack of contamination on the spacecraft windows; and the continuous assistance and evaluation of the science support room at the Manned Spacecraft Center.


The crew also made an inertial measurement unit alignment with a sighting of the planet Jupiter (the first time a planet had been used) and performed a number of daylight star sightings, landmark sightings, and star sextant sightings. During two successive revolutions, at 192:43 and 194:13, the crew successfully tracked the Pegasus III satellite at a range of 1,000 n mi. Pegasus III had been launched on 30 July 1965.


While over Hawaii, the crew made a sighting of the ascent stage from 222:38:40 to 222:45:40.


The service propulsion system had been fired for the seventh time at 169:30:00.36, a 24.90-second maneuver that raised the orbit to 253.2 by 100.7 n mi and established the desired conditions for the nominal deorbit point. If the service propulsion system had failed at deorbit, the reaction control system could have conducted a deorbit maneuver from this apogee condition and still landed near the primary recovery area. The deorbit maneuver was accomplished after 151 orbits with the eighth service propulsion firing, an 11.74-second maneuver at 240:31:14.84. It was performed one revolution later than planned because of unfavorable weather in the planned recovery area.





The service module was jettisoned at 240:36:03.8, and the CM entry followed a primary guidance system profile. The command module reentered Earth’s atmosphere (400,000 feet altitude) at 240:44:10.2 at a velocity of 25,894 ft/sec. Although the service module could not survive entry intact, radar tracking data predicted impact in the Atlantic Ocean at a point estimated to be latitude 22.0° north and longitude 65.3° west, 175 n mi downrange from the CM.


The parachute system effected splashdown of the CM in the Atlantic Ocean at 17:00:54 GMT (12:00:54 p.m. EST) on 13 March. Mission duration was 241:00:54. The impact point was about 2.7 n mi from the target point and 3 n mi from the recovery ship U.S.S. Guadalcanal. The splashdown site was estimated to be latitude 23.22° north and longitude 67.98° west. After splashdown, the CM assumed an apex-up flotation attitude. The crew was retrieved by helicopter and was aboard the recovery ship 49 minutes after splashdown. The CM was recovered 83 minutes later. The estimated CM weight at splashdown was 11,094 pounds, and the estimated distance traveled for the mission was 3,664,820 n mi.


At CM retrieval, the weather recorded onboard the Guadalcanal showed scattered clouds at 2,000 feet and broken clouds at 9,000, visibility 10 n mi, wind speed 9 knots from 200° true north, air temperature 79° F, and water temperature 76° F, with waves to seven feet from 340° from true north.


The crew left the Guadalcanal by helicopter at 15:00 GMT on 14 March and arrived at Eleuthera, Bahamas, at 16:30 GMT. From there, they were flown to Houston.


The CM was offloaded from the Guadalcanal on 16 March at the Norfolk Naval Air Station, Norfolk, Virginia, and the Landing Safing Team began the evaluation and deactivation procedures at 16:00 GMT. Deactivation was completed on 19 March. The CM was then flown to Long Beach, California, and trucked to the North American Rockwell Space Division facility at Downey, California, for postflight analysis, where it arrived on 21 March.





The following conclusions were made from an analysis of post-mission data:


  1. The onboard rendezvous equipment and procedures in both spacecraft provided the required precision for rendezvous operations to be conducted during a lunar landing mission. The CSM computations and preparations for mirror-image maneuvers were completed on time by the command module pilot.


  1. The functional operation of the docking process of the two spacecraft was demonstrated. However, the necessity for proper lighting conditions for the docking alignment aids was illustrated.


  1. The performance of all systems in the extravehicular mobility unit was excellent throughout the entire extravehicular operation. The results of this mission, plus satisfactory results from additional qualification tests of minor design changes, provided verification of the operation of the extravehicular mobility unit on the lunar surface.


  1. The extent of the extravehicular activity indicated the practicality of extravehicular crew transfer in the event of a contingency. Cabin depressurization and normal repressurization were demonstrated in both spacecraft.


  1. Performance of the lunar module systems demonstrated the operational capability to conduct a lunar mission, except for the steerable antenna, which was not operated, and the landing radar, which could not be fully evaluated in Earth orbit. None of the anomalies adversely affected the mission. The concepts and operational functioning of the crew/spacecraft interfaces, including procedures, provisioning, restraints, displays, and controls, were satisfactory for manned lunar module functions. The interfaces between the two spacecraft, while both docked and undocked, were also verified.


  1. The lunar module consumable expenditures were well within predicted values, thus demonstrating adequate margins to perform the lunar mission.


  1. Gas in the CM potable water supply interfered with proper food rehydration and therefore had some effect on food taste and palatability. Lunar module water was acceptable.


  1. Orbital navigation of the CSM, using the yaw-control technique for landmark tracking, was demonstrated and reported to be adequate. The star visibility threshold of the CM scanning telescope was not definitely established for the docked configuration; therefore, platform orientation using the sun, the Moon, and planets may be required if inertial reference is inadvertently lost during translunar flight.


  1. Mission support, including the Manned Space Flight Network, adequately provided simultaneous ground control of two manned spacecraft.