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Little Joe II

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For the Mercury program flight, see Little Joe.

The Little Joe II tests were tests for the LES.

Little Joe II
Apollo spacecraft CSM-002 atop a Little Joe II launch vehicle at White Sands Missile Range. (NASA)
Stages 2
0 - Booster Engines 5 * Recruit rockets
Thrust 167 kN x 5 = 836 kN
Burn time 1.53 seconds
Fuels Solid
1 - Sustainer Engines 2 * Algol rockets
Thrust 465 kN x 2 = 930 kN
Burn time 40 seconds
Fuels Solid
2 - Second stage Engines 2 * Algol rockets
Thrust 465 kN x 2 = 930 kN
Burn time 40 seconds
Fuels Solid
A-004 version Launch January 20, 1966
Payload 30,000 lb (14,000 kg)

Little Joe II program

From August 1963 to January 1966, a series of unmanned flight tests were conducted at the White Sands Missile Range to demonstrate the adequacy of the Apollo launch escape system and to verify the performance of the command module earth landing system. The launch vehicle used for five of these tests was the Little Joe II. Its predecessor, the Little Joe, had been used in testing the launch escape system for the Mercury spacecraft. In addition to the Little Joe II flights, two pad abort tests were conducted in which the launch escape system was activated at ground level.

The program was originally planned to be conducted at the U.S. Air Force Eastern Test Range at Cape Kennedy. However, because of a heavy schedule of high-priority launches at that facility, other possible launch sites were evaluated. Launch Complex 36 at White Sands Missile Range, previously used for Redstone missile test, was ultimately selected as the most suitable for meeting schedule and support requirements. Also, the White Sands Range allowed land recovery which was less costly and complicated than the water recovery procedure that would have been required at the Eastern Test Range or at the NASA Wallops Island facility.

The program was conducted under the direction of the Manned Spacecraft Center (now Johnson Space Center), Houston, TX, with joint participation by the prime contractors for the launch vehicle (General Dynamics/Convair) and spacecraft (North American Rockwell). The White Sands Missile Range administrative, range, and technical organizations provided the facilities, resources, and services required. These included range safety, radar and camera tracking, command transmission, real-time data displays, photography, telemetry data acquisition, data reduction, and recovery operations.

Launch vehicle development

Man-rating of the launch escape system was planned to be accomplished at minimum cost early in the Apollo program. Since there were no reasonably priced launch vehicles with the payload capability and thrust versatility that could meet the requirements of the planned tests, a contract was awarded for the development and construction of a specialized launch vehicle. Fabrication of the detail parts for the first vehicle started in August 1962, and the final factory systems checkout was completed in July 1963. There was an original fixed-fin configuration and a later version using flight controls.

Four Apollo rocket assemblies, drawn to scale: Little Joe II, Saturn I, Saturn IB, and Saturn V.

The vehicle was sized to match the diameter of the Apollo spacecraft service module and to suit the length of the Algol rocket motors. Aerodynamic fins were sized to assure that the vehicle was inherently stable. The structural design was based on a gross weight of 220,000 pounds (100,000 kg), of which 80,000 pounds (36,000 kg) was payload. The structure was also designed for sequential firing with a possible 10-second overlap of four first-stage and three second-stage sustainer motors. Sustainer thrust was provided by Algol solid-propellant motors. Versatility of performance was achieved by varying the number and firing sequence of the primary motors (capability of up to seven) required to perform the mission. Recruit rocket motors were used for booster motors as required to supplement lift-off thrust.

A simplified design, tooling, and manufacturing concept was used to limit the number of vehicle components, reduce construction time, and hold vehicle cost to a minimum. Because overall weight was not a limiting factor in the design, over designing of primary structural members greatly reduced the number and complexity of structural proof tests. Whenever possible, vehicle systems were designed to use readily available off-the-shelf components that had proven reliability from use in other aerospace programs, and this further reduced overall costs by minimizing the amount of qualification testing required.

The Little Joe II launch vehicle proved to be very acceptable for use in this program. Two difficulties were experienced. The qualification test vehicle (QTV) did not destruct when commanded to do so because improperly installed primacord did not propagate the initial detonation to the shaped charges on the Algol engine case. The fourth mission (A-003) launch vehicle became uncontrolled about 2.5 seconds after lift-off when an aerodynamic fin moved to a hard over position as the result of an electronic failure. These problems were corrected and the abort test program was completed.

Minor spacecraft design deficiencies in the parachute reefing cutters, the drogue and main parachute deployment mortar mountings, and the command module/service module umbilical cutters were found and corrected before the manned Apollo flights began. However, all command modules flown achieved satisfactory landing conditions and confirmed that, had they been manned spacecraft, the crew would have survived the abort conditions.

Launch configuration summary

Item QTV A-001 A-002 A-003 A-004
Launch weight 25,930 kg 26,281 kg 42,788 kg 80,372 kg 63,381 kg
Payload 10,988 kg 11,492 kg 12,561 kg 12,626 kg 14,717 kg
Airframe weight 14,942 kg 14,785 kg 29,320 kg 67,745 kg 48,623 kg
Liftoff thrust 49 kN 49 kN 1,600 kN 1,395 kN 1,766 kN
Fin – fixed/control F F C C C
1st stage Recruit 6 6 4 0 5
1st stage Algol 1 1 2 3 2
2nd stage Algol 0 0 0 3 2



Surviving examples

Specifications

  • Little Joe II
    • Thrust: 49 to 1,766 kN
    • Length: 10.1 m without - CM/SM/LES
    • Length: 26.2 m with CM/SM/LES
    • Diameter: 3.9 m body
    • Fin span: 8.7 m
    • Weight: 25,900 to 80,300 kg
    • Fuel: solid
    • Burn time: ~50 s


  • Algol rocket
    • Thrust: 465 kN each
    • Length: 9.1 m
    • Diameter: 1 m
    • Weight full: 10,180 kg
    • Weight empty: 1,900 kg
    • Fuel: solid
    • Burn time: 40 s


  • Recruit rocket (Thiokol XM19)
    • Thrust: 167 kN
    • Length: 2.7 m
    • Diameter: 0.23 m
    • Weight: 159 kg
    • Fuel: solid
    • Burn time: 1.53 s


References

  1. ^ Alamogordo's Space Center Retrieved: 14 June 2008.


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