Self-Unloading, Unmanned, Reusable Lunar Lander Project - Abstract
Self-Unloading, Unmanned, Reusable Lunar Lander Project
Kevin Cowan, et. all
May 5, 1991
The future of the U.S. space program outlined by President Bush calls for a permanently manned lunar base. A payload delivery system will be required to support the buildup and operation of that lunar base. In response to this goal, RS Landers has developed a conceptual design of a self-unloading, unmanned, reusable lunar lander. The lander will deliver a 7000 kg payload, with the same dimensions as a space station logistics module, from low lunar orbit (LLO) to any location on the surface of the moon.
This executive summary briefly introduces the technical aspects of the design as well as the management structure and project cost.
This concept is a product of rigorous brainstorming followed by meticulous inspection of the resulting ideas. The payload delivery system consists of a lander, unloader, and payload.
As the figure shows, the payload and the unloader are loaded in an inverted position on top of the lander. After post-landing stabilization on the lunar surface, the entire structure will rotate 180ˇ with respect to the legs. This rotation will take at least thirty minutes.
When the rotation is complete, the unloader will be lowered to the surface. The unloader will then drive out between the legs and deliver the payload to its desired location. The unloader has a range of 5 km when loaded with the payload. The 5 km range was needed because it was determined that the lander should land at least 1-2 km away from the lunar base. This is necessary in order to avoid excessive plume damage. Once the payload has been delivered, the unloader can return to low lunar orbit (LLO) with the lander, or it can remain on the surface to await the lander's return.
Solid core nuclear motors were chosen for use on the lunar lander. These motors have an optimistically projected specific impulse of 1200 sec. and thrust to weight ratio of 11.3. The maximum required thrust occurs during the descent phase of the mission, and it is 22,584 lbf.
It is not currently known whether a three motor configuration or a single motor configuration would be superior for use on the lander. For conventional motors, the three motor configuration is recommended for situations of engine out. There are studies being done to determine the effect of clustering nuclear motors. It may be necessary to use one nuclear motor with redundant turbopumps. All lander drawings in this document, however, show the three motor configuration.
The detailed design of the mechanical components of the various payload unloading mechanisms is beyond the scope of this study, however, there are a few points that should be considered in their design. The areas that were considered are the types of electric motors, bearings, and drive train or gear reduction system that should be used.
The motors that are most promising for the RS Landers La Rotisserie concept use direct current, deliver moderate torque, medium rotation rates (around 1000 rpm), and are of a brushless design.
Coated bearings are suggested for use on the lunar lander. Lubricants will prove to be ineffective in the harsh lunar environment. They will either become filled with dust, freeze up, or boil off. Possible bearing coatings include Teflon¨, Nomex¨, and diamond. Diamond coatings can be applied using Chemical Vapor Deposition.
A harmonic drive system is recommended for use on the lunar lander. Harmonic drives have fewer moving parts than the conventional gear box. They are therefore less massive. Harmonic drive systems use flexible splines that wear faster than conventional gear box components, but with the advent of advanced materials, the harmonic drive can be designed to more than meet the lander's requirements.
The lander trajectories have been designed and optimized using a computer program called Lander. Program Lander was developed by Eagle Engineering in Houston, TX to simulate the ascent and descent phases of a lunar landing mission.
The landing site location of the Apollo 15 mission was chosen for the lunar lander simulation. The resulting total delta-velocities were 1.839 km/sec for ascent and 1.92 km/sec for descent. The flight times were 50 minutes for ascent and 63.25 minutes for descent. The use of the solid core thermal nuclear propulsion system greatly increased the efficiencies of the trajectories.
Structures and Materials
The lander structure provides connectivity and integrity to all of the lander's systems. The central box of the lander structure carries all of the loads generated by the subsystems. This box is a truss structure enclosed by honeycomb core panels. The truss structure is strong enough to support the loads generated by the subsystems, and its lightweight panels protect the subsystems from solar radiation, dust, and micrometeorites.
The landing gear is composed of four struts that are lightweight planar trusses with Apollo style Lunar Module landing pads. To enable the lander to remain level on an incline of up to 8ˇ, a terrain adaptive system is incorporated into the landing gear.
Aluminum-lithium alloys were chosen as the main construction material for the lander. In addition to having the advantages of standard aluminum alloys, they can have a high tensile strength (over 100 ksi), along with increased weldability and a higher cryogenic strength.
Guidance, Navigation, & Control
The purpose of the guidance, navigation, and control (GNC) system is to determine the linear and angular position, velocity, and acceleration of the lander, to compare that data with the desired state, and to make corrections when necessary. The desired state of the lander will be provided by the predetermined trajectory analysis for each specific mission.
The GNC system is made up of three components: Sensors, Computer, and Control. The sensors provide information on the state of the lander. The computer evaluates the data from the sensors and instructs the control mechanisms. The control mechanisms then change the state of the lander by their action.
Three levels of sensors are used for redundancy. During optimum operating conditions, several components of each level of redundancy will be used. The primary, secondary, and emergency sensor arrays rely on a radar imaging/altimeter system, several sets of accelerometers and gyros, a transponder system, a close proximity altitude detection device, and the communications system. The communications system is only used as a sensor for emergency situations.
The onboard navigation computer will be a fault tolerant advanced computer that will be capable of out-performing today's most advanced Cray computer. The rapid pace of computer and software development has shown that a system of this type will not only be possible, but will have little mass and power consumption. The navigation computer will be responsible for monitoring the output and status of each sensor, monitoring the status of and providing input for each of the control devices, and providing an interface between the two.
The lander will utilize three control techniques: momentum exchange devices, small directional thrusters, and gimbaled/throttled main engines. While some redundancy exists using all three systems, the optimum operating conditions will use each technique where best suited.
The control of the unloader will be primarily automated with a remote control system as a backup. The unloader will have optical sensors which will inform the unloader's onboard computer of obstacles. The computer will then instruct the wheel motors to make the required adjustments. The unloader will be in constant communication with the lander, in case it becomes necessary to employ the back up remote control system.
The communications systems provide three basic functions: telemetry, command, and tracking. There are three areas of communications that will be performed by the lunar lander system: Lander/Earth, Lander/OTV, and Lander/Unloader.
S-band (2.3 GHz) will be used for direct communications between the lander and earth. The antenna on the lander will be a parabolic dish with steering capabilities similar to that on the Apollo spacecraft. The Apollo pointing system is more than sufficient for the communication link with earth.
It is also recommended that a communications satellite for Earth link capabilities be placed in a halo orbit about the L2 Lagrangian point. The satellite would allow transmissions to be made between the lander and earth when the lander is on the far side of the moon.
Communications between the OTV and the lander will be done with a VHF system. This system will be necessary for docking procedures. Once the lander is docked with the OTV a data feed umbilical will be connected to the lander by means of a manipulator arm on the OTV.
Communications between the lander and unloader will be done using a UHF system. The UHF receivers and transmitters are small, lightweight, and require little power. The UHF antennas are also small.
The energy for the power system is provided by the heat generated during engine cool down cycles. A power conversion loop transforms the heat into electrical energy, which is then stored in rechargeable Na-S batteries on the lander and the unloader. The conversion loop also serves to cool down the nuclear motors and keep the batteries at a higher operating temperature.
Two sets of batteries provide 11 kWh of power on both the lander and the unloader. The power for the unloader allows it to carry the payload 5 km at a speed of 2.5 km/hr. In the event that the unloader remains on the surface for an extended period, two solar arrays totalling 20 sq. m. are mounted on the unloader. These GaAs/Ge arrays are able to recharge the batteries fully in about one solar day. The solar arrays are also used to heat the unloader's batteries during payload transport.
Thermal control will be done using several methods. The first method will employ the use of a cryogenic refrigeration system that will be powered by the power generation loop. The second method will employ the use of 2 1/2" of multi-layer insulation on the propellant tanks and other areas that require thermal control. Heat exchangers on the power generation loop will also be used to keep certain areas of the lander warm.
The final method that will be used is the use of two radiation/thermal protection umbrellas. These umbrellas will be deployed from the landing struts after the complete rotation sequence has been performed. The umbrellas will help to reduce the work load on the refrigeration system considerably.
The management structure of RS Landers was designed for speed of communications between all group members. Progress and problems, for example, are reported directly to the C.E.O. and the Technical Supervisor. The final design is the compilation of input from all group members. All milestones for the project were completed on time.
The actual total cost of the project was $27,457. This is 2% less than the cost that was estimated at the start of the project.
CSR/TSGC Team Web