Little Joe II
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For the Mercury program flight, see Little Joe 2.
| Little Joe II | ||
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| Missing image Little_Joe_II_A-004.jpg 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) | |
| Contents |
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 vehicle 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.
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 |
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
External Links
- The Apollo Spacecraft: A Chronology (http://www.hq.nasa.gov/office/pao/History/SP-4009/cover.htm)
- Apollo Program Summary Report (http://history.nasa.gov/apsr/apsr.htm)
- Little Joe II Qualification Test Flight Report - September 1963 (PDF) (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19790076764_1979076764.pdf)
- Little Joe II Progress Report - December 1964 (PDF) (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19780070078_1978070078.pdf)
- Little Joe II Mission A-003 - April 1965 (PDF) (http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19790078425_1979078425.pdf)
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