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BALLSAR Project: (Ballistic Arbitrary Location Lunar Sampler and Retriever)

Greg Moroney, et all

December 7, 1999

Executive Summary

1.0 Purpose and Scope

The Fall 1999 mission design team, Ballistic Arbitrary Location Lunar Sampler and Retriever (BALLSAR), proposes a lunar sampler which will collect and analyze 1 kg samples from the lunar surface. In order to warrant a mission for lunar samples, there need to be several justified reasons for lunar samples. Some reasons are to determine the history of the moon and possibly the Earth, determine the composition of the moon for possible future lunar bases, fill a need of lunar rocks for research, and for commercial sales of lunar rocks.

2.0 Heritage Missions

Several missions were studied as the BALLSAR mission was designed, including Apollo, Clementine, Lunar Prospector, and Lunar Retriever 1. The Apollo missions were used for ascent and descent delta v's. Clementine and Lunar Prospector had similar communications systems, and the Lunar Retriever assisted in some phases of the mission.

3.0 Launch

The preliminary choice for the launch vehicle is the Ariane 5. This vehicle was chosen because it is large enough so that the sampler will be able to share the launch with SELENI. Sharing the launch will lower the cost of the mission.

4.0 Trajectories

The BALLSAR mission will use a Hohmann transfer trajectory to arrive at the moon. The preferable trajectory sequence will begin with the lunar sampler aboard a launch vehicle. An upper stage of the launch vehicle will put the sampler and its accompanying boosters into an elliptical LEO to GEO transfer orbit. At perigee, the sampler will be released where it will begin its lunar transfer orbit using the first stage booster. This sequence is preferable because the trajectory requires the least amount of Dv. Hohmann transfer Dv figures were calculated for all the needed trajectories, as a first look at the minimum Dv needed to accomplish the mission. Apollo Dv figures were used when appropriate instead of Hohmann Dv figures. After the sampler has conducted its mission, the ascent stage will dock with the orbiting second stage. The second stage will be used to return the sample capsule back to Earth.

5.0 Mass and Fuel Estimates

A numerical routine was written in MATLAB to calculate approximate structural masses for each of the mission components and fuel masses for each of the burns. The approximate fuel mass for the first booster stage was 788 kg and the wet mass of that stage and the lander is 1911 kg. The fuel required in the lander is 628 kg and the total wet mass of the lander is approximately 829 kg. The second booster stage only requires a fuel mass of 76 kg and has a wet mass of 238 kg.

6.0 Subsystems

6.1 Propulsion

Four clusters of four reaction thrusters will be used for attitude control on both the second stage booster (orbiter) and the lander. A STAR-37F solid rocket engine will be used for the trans-lunar trajectory. A Marquardt R-40 engine will be used for the main engine of the lander. A Marquardt R-4D engine will be used as the engine for the ascent stage of the lander. A Thiokol STAR-17A engine will be used for the second stage return booster.

6.2 Guidance, Navigation and Control

A GN&C system will be used on both the orbiter and the lander. Both vehicles will be 3-axis stabilized and will use both sun and horizon sensors to determine attitude in all three axes. The lander will use four sets of four small thrusters for attitude corrections while maneuvering on the lunar surface. The orbiter will use four reaction wheels to control attitude and four sets of four small thrusters for momentum dumping. The lander will use radar for decent rates, imaging, and automated docking.

6.3 Communications

A communications system is needed to link the orbiter and the lander to the Earth. Each vehicle will have a direct line of communication to Earth, and a two-way coherent mode will be used. Both the orbiter and the lander will have a high gain parabolic dish, an antenna, and transponders. The difference in communication systems between the lander and the orbiter will be that the orbiter will have a solid state recorder and the lander will not need one because it will constantly relaying data to Earth.

6.4 Command and Data Handling

The C&DH subsystem on the orbiter will be a typical system and will be 7500 cm3, weigh 5.5 kg, and require 15 W of power. The C&DH on the lander will be a complex system because of the need for a computer interface. This system will be 13000 cm3, weigh 10 kg, and use 20 W of power.

6.5 Power

Once on the moon, the lander and the orbiter will be powered by GaAs solar panels with backup provided by Nickel Hydrogen batteries. The solar arrays used on the orbiter will need to be 0.706 m2 and on the lander they will need to be 0.447 m2.

6.6 Thermal

Both the orbiter and the lander will require ground spacecraft equipment heaters to provide warmth to the subsystems contained in the structures. Multi Layer Insulation blankets will also be used to provide warmth to the subsystems during cold periods. Fins will release heat to regulate the subsystem temperatures. A thermal shield consisting of ceramic tiles will be used for capsule re-entry.

6.7 Lander

The vehicle that will be used to acquire the lunar samples is loosely based on the Delta Clipper and Clipper Graham vehicles. The descent stage of the lander will perform a deorbit burn from lunar orbit and land vertically on the moon. The sample method design involves a robotic arm with a scoop attachment, which will collect and sift the samples. The scoop will also scrape away the surface of the lunar soil to minimize sample contamination. To move from one sample site to another, an impulse burn will be used to put the lander on a ballistic trajectory. The ascent stage of the lander will place the sample capsule in an orbit, which will autonomously intercept and dock with the orbiting booster. When docked, the second stage of the lander will be jettisoned and the sample capsule and second stage of the booster will return to Earth. Once near Earth orbit, the capsule will separate from the booster, re-enter and parachute to Earth. When the lander has landed, a GPS signal will be activated for quick retrieval. The samples will then be analyzed.

7.0 Group Organization

Christopher Garland is the team leader of the BALLSAR mission team, as well as the project head for the trajectory and mass estimation teams. Jennifer Sonderfan is the project head for the launch vehicle and the control systems teams. Greg Moroney leads the propulsion, lander operations, and return teams.

8.0 Cost Analysis

The cost of the mission design review conducted by BALLSAR was $9950. The actual mission costs have yet to be determined.

9.0 Project Deliverable and Schedule

The final presentation will be delivered the week of November 29, 1999. The final design review report, model, project book and poster will also be presented by December 7, 1999.

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