Propulsion Subsystem - Lamar
INTERDISCIPLINARY SPACECRAFT DESIGN PROJECT
LAMAR - PROPULSION TEAM
CONCEPTUAL DESIGN REPORT
ME-431 Integrated Systems Design
Dr. W. E. Simon
October 13, 1995
TABLE OF CONTENTS
|Launch Propulsion System||1|
|Main Propulsion System||3|
|1st Conceptual Design||5|
|2nd Conceptual Design||6|
|3rd Conceptual Design||6|
|4th Conceptual Design||6|
|Reaction Control System||7|
|1st Conceptual Design||8|
|2nd Conceptual Design||8|
Propulsion Group Conceptual Design Report
One of the most important resources necessary for human presence on the Moon is oxygen. Demonstration of oxygen production on the moon would be a major stepping stone toward the establishment of a permanent lunar base. The Texas Space Grant Consortium is sponsoring a space related student design project during the 1995-96 school year. The goal of this project is to provide the students from each of the participating TSGC institutions with a meaningful and exciting design experience. These teams of students and faculty from the Texas Space Grant Consortium academic institutions shall develop a preliminary design for a small lunar lander to demonstrate production of oxygen on the surface of the Moon using in situ resources. The team from each academic institution will chose one or more spacecraft subsystems and will develop a design for that subsystem. The final design proposal will be submitted to NASA under a small spacecraft program with the intent of partnering with Texas industries to build and fly the proposed mission.
Lamar University has four separate teams associated with this project. This report is in compliance with the requirements of the Interdisciplinary Spacecraft Design Project by the Lamar Propulsion Team.
The propulsion group will have three main responsibilities: launch vehicle selection; main propulsion system (MPS) design; and reaction control system (RCS) design. The launch vehicle will primarily be selected according to the total spacecraft mass, but will restrict the structural design due to payload fairing restrictions (length and principal diameter). The purpose of the main propulsion system is to supply sufficient thrust at proper times and locations in order to fly the spacecraft from the earth orbit to the lunar orbit and then to the lunar surface. Attitude control, or maintaining of the orientation, is served partially through the Reaction Control System. In addition to these selections and designs, the group will need to design all supporting propulsion subsystems and integrate the propulsion system into the spacecraft.
Launcher Propulsion System
The main purpose of the launch propulsion system is to lift the launch vehicle and its payload from the launch pad and place the payload into an Earth orbit. This orbit may be a high or low earth orbit. This will depend upon the launch vehicle selected. A larger launcher has the capability of achieving a higher energy orbit. This creates a trade-off. With a large launch vehicle, a smaller Main Propulsion System may be used during the flight. This is due to the fact that a smaller energy would be required of the MPS. Conversely, a smaller launch vehicle would lead to a larger MPS. An upper stage may be used, such as a PAM-D, to get the spacecraft from a lower to a higher orbit.
The term payload includes all hardware above the launch vehicle to spacecraft interface. The launch vehicle payload consists of the entire spacecraft and the booster adapter. Launch vehicle selection depends on the size and mass of the final spacecraft. The mass of the vehicle for this project has been projected to be between 500 and 800 kg. Therefore, a range of approximately 500 - 1000 kg. will be used for the launch vehicle selection.
The following is a list of conceptual design concepts for this mission.
The DELTA II is a multi-stage launch vehicle which would insert the spacecraft into earth orbit. The first stage of the DELTA II is the main engine which supplies lift-off and enters the spacecraft into flight. The second stage is a propulsion system which contains guidance and control equipment that provides guidance sequencing and stabilization signals for both first and second stages. The third and final stage of the DELTA II jettisons the remainder of the launch vehicle leaving only the spacecraft and its propulsion systems.
There are several reasons to choose the DELTA II. It provides more energy for lift-off (nine boosters, six of which ignite at lift-off and stage at 57 seconds). The first and second stages use inertial guidance and provide control moments from gimbaled engines. The payload mass of the DELTA II is 800 kg. This is sufficient since the payload mass requirement has been approximated at 500 to 1000 kg. The allowable principal diameter is 2.5 m., and the allowable length is 2.8 m. These parameters would govern the size of the structure of the spacecraft. Another reason to choose the Delta II launch vehicle is its high reliability.
Lockheed Launch Vehicle (LLV)
The LLV is also a multi-stage launch vehicle that would insert the spacecraft into earth orbit. The first stage of the LLV provides lift-off and enters the spacecraft into flight until ignition burn out. The second stage provides propulsion during the first stage separation and through ignition burn out. The third stage serves the same purpose as the second stage.
Benefits of the LLV include the use of Thiokol solid rocket boosters for better performance. The 800 kg. payload mass meets the requirements for this mission. One handicap of choosing the LLV is that it is still in development. This creates a technological risk.
OSC Pegasus Vehicle
The Pegasus serves the same purpose as both the DELTA II and the LLV. Again,
the first stage provides lift off and flight entrance. Aerodynamic fins are used for direction control. The second and third stages contain vectorable nozzles that control the stages.
A benefit of the OSC Pegasus is that it employs the B-52 as a launch platform. This reduces the propellant needed to achieve orbit and provides variable launch azimuths and locations to place the spacecraft in different orbit inclinations. A drawback, however, is that it can put only very small payloads into orbit.
The space shuttle consists of a reusable delta-winged space plan called an orbiter and two solid propellant rocket boosters, which are recovered and reused. The Shuttle uses two solid-rocket boosters and three liquid-oxygen and liquid-hydrogen engines for propulsion.
Benefits include a large payload bay. On the down side, it is relatively expensive to put a payload on the shuttle due to its cost of integration and fail-safe mechanisms for the manned spacecraft. There is also a backlog of missions waiting to use the Space Shuttle.
Main Propulsion System
The Main Propulsion System (MPS) is responsible for transporting the lunar lander from earth orbit to lunar orbit and possibly aiding in the descent phase to the surface of the moon. The MPS must provide the substantial DETA-V's throughout the lunar mission. The DETA-V's are the changes in velocity needed to move the lander through each mission phase. These DETA-V's are a result of the Navigation, Guidance and Control preliminary trajectory calculations. Using these calculations, conceptual designs can be made considering certain requirements and assumptions.
The MPS 's specific design responsibilities take over when the spacecraft separates from the launch vehicle. After separation the lander travels through several phases each requiring a certain DETA-V. These stages are Trans-Lunar Injection, Lunar Orbit Insertion and Descent phase.
The first DETA-V is called the Trans-Lunar Injection burn, (TLI). This phase starts at a low earth orbit. A certain DETA-V is required to increase from the energy of a circular orbit to a higher energy elliptical orbit called the transfer orbit. In the transfer orbit, the perigee is the altitude of the low earth orbit and the apogee is at the altitude of the moon's orbit. At the apogee, the spacecraft has a velocity that is lower than the velocity of the moon due to the moon being in a higher energy circular orbit. The differences in velocities is called V infinity minus. At this point, the spacecraft is influenced more by the gravity field of the moon than that of the earth. The spacecraft is then subjected to a hyperbolic trajectory.
Lunar Orbit Insertion
Once the spacecraft approaches lunar orbit, the second DETA-V is performed in opposite direction of the velocity. This is done to decrease the energy of the hyperbolic orbit to the energy of a circular orbit around the moon.
The Descent Phase is the final mission phase of the MPS. This trajectory includes the De-Orbit Burn which decreases the energy of the lunar orbit enough that the new orbit intersects the surface of the moon. Then a deceleration burn is required to slow the spacecraft from orbital velocity to zero velocity at the surface of the moon.
In the conceptual design phase the MPS must consider several requirements. These requirements are restartability, throttleability, thrust for sustained hover, reliability and low cost. To minimize technological risks and development costs, an existing propulsion system is strongly considered for use. Then, the different combinations of thrusters and engines can be determined for this mission. The MPS design should also include engine selection, fuel and oxidizer tank sizing, fuel and oxidizer pressurization system design and fuel and oxidizer delivery system design. This information is then passed on to the other groups in the project.
Trajectory estimates, i.e. DETA-V's, from the Navigation, Guidance and Control Group are needed to begin the MPS design. The values of the DETA-V's are based on the assumption that the spacecraft has attained a 296 km. lower earth orbit. A safety factor of 20% is included in the totals to account for the possibility of unfavorable launch conditions.
|Low Earth Orbit Corrections||0||0||10|
|Earth Departure Burn||0||0||3108|
|Trajectory Correction Maneuver 1||0||0||20|
|Lunar Orbit Insertion||0||0||825|
|+20% Safety Factor||3||3||1150|
Conceptual Designs For The Main Propulsion System (MPS)
The first conceptual design for the MPS has three liquid fuel engines, each supplying propulsion for a different phase of the mission. The first stage engine shall provide the required DETA-V that will complete the Trans-Lunar Injection Phase. It will then be staged from the spacecraft. The remainder of the MPS will consist of three auxiliary thrusters, that delivers the specified DETA-V that completes the Lunar Orbit Insertion Phase. These thrusters shall also be used for the Descent Phase, and share responsibility with the RCS.
- Simplistic design
- More control of the spacecraft
- Takes load off the RCS
- More mass carried to the lunar surface
- Possible high cost
The second conceptual design for the MPS has three main engines that allows the lunar lander to complete each specified mission phase. The first engine shall use solid fuel to obtain the specified DETA-V needed complete the Trans-Lunar Injection Phase whereby, it will stage from the spacecraft. The second engine will use solid fuel to generated the DETA-V needed to complete the Lunar Orbit Insertion Phase, thereafter staging form the spacecraft. The final engine shall use liquid fuel to complete the Descent Phase of this mission, and also provide assistance to the RCS.
- 1. Less cost to fuel
- Unwanted mass is discarded
- Lighter load
- More complicated design
The third conceptual design for the MPS has a two main engines that allows the lunar lander to complete each specified mission phase. The first engine uses solid fuel to generate the required DETA-V that will complete the Trans-Lunar Injection Phase, whereby it will be staged from the spacecraft. The remaining MPS will consist of a solid fuel engine, which will be used to complete the Lunar Insertion Phase, and the Descent Phase, that includes the De-Orbit Burn. Once the Lunar Insertion and Descent Phases have been completed this solid fuel engine shall be staged from the spacecraft. Thus, remain three auxiliary thrusters that will complete the Descent Phase, and possibly assist the RCS.
- Unwanted mass discarded.
- Costly high-tech design
This fourth conceptual design for the MPS consist of one main liquid engine, three small liquid engines, and three auxiliary thrusters. The main engine is responsible for carrying the spacecraft from the Trans-Lunar Injection to the De-Orbit Burn Phase, whereby it will be staged. The remaining MPS attached to the spacecraft will consist of three small liquid engines that are used to complete the Descent Phase of this mission. Thus, leaving only three thrusters that assist the RCS for the Deceleration Phase.
- More controllable
- Complex design
- High cost
Reaction Control System
The propulsion group will provide a theory of operation for the Reaction Control System and all subsystems. The RCS design should include the thruster selection, thruster positioning, fuel tank sizing, fuel pressurization system design, and fuel delivery system design. The Propulsion Specifications require the RCS to use monopropellant thrusters due to their simplicity. It also requests coupled thrusters whenever possible.
The main purpose of the Reaction Control System is to provide coarse attitude control throughout the flight, including the descent to the surface of the moon. The RCS is used to correct the spacecraft orientation whenever needed throughout the flight of the spacecraft from earth orbit to lunar orbit and to the lunar surface.
The Reaction Control System must maintain this stabilization. Both slight and severe attitude corrections will be necessary. The RCS must be prepared to handle both. During burn maneuvers, larger corrections could be required to keep the spacecraft oriented properly. The spacecraft will be unstable when the main propulsion system burns. For this reason the reaction control system must act quickly, and forcefully. The RCS thrusters are typically used as impulsive maneuvering jets. One pulse gets the spacecraft moving in one direction, and another pulse stops that motion. During these main engine burns it may be required that the thrusters pulse very frequently. However, during the majority of the trip the thrusters may pulse very infrequently.
The event of slight coarse corrections may arise due to overall design of the spacecraft. The center of mass may not be along the axes of the spacecraft, causing an imbalance in the structure. That imbalance would cause a moment, resulting in a rotation of the craft that must be corrected. The RCS must counteract severe rotation rates and may be called upon to implement small attitude corrections. Due to the varying requirements and circumstances, a flexible attitude control system is essential.
During the descent phase all of these possibilities must be guarded against, as well. The spacecraft should remain vertical and its flight should be perpendicular to the surface. The spacecraft should remain oriented along its velocity vector during the descent, except for controlled descent trajectory corrections. The RCS should keep the spacecraft oriented vertically so that the spacecraft will be prepared for the landing. During landing there will be a site selection process that will require the spacecraft to avoid landing areas that are not level enough upon which to land. This type of translational trajectory correction could also be implemented by the RCS.
The MPS supplies the propulsion necessary to propel the spacecraft from the earth orbit to lunar orbit, and it then provides propulsion in descent to ensure that the craft will have a velocity as close to zero as possible. Therefore, the RCS will coincide with the MPS, and will the two systems will have to be compatible with one another. An assumption being made is that the MPS will power the flight and the landing, and that the RCS will simply provide attitude control during both phases.
The spacecraft will contain momentum wheels within the structure to provide for fine attitude control. The momentum wheels rotate at a constant rate until a moment I applied. Slowing down the wheels, or increasing their speed will create a moment in the opposite direction of the acting moment. This could be used for attitude corrections that are associated with slower rotational rates. However, the momentum wheel can only provide as large a moment as its initial and maximum speed will allow. When one comes to a complete stop, or reaches its maximum speed, it can no longer produce an additional moment. For faster rates of rotation, or to bring the momentum wheel back to its nominal speed, thrusters may be used. The following conceptual designs assume a hexagonal-shaped lander.
The first conceptual design being submitted for the RCS consists of a spherical assembly on the top and bottom of the lander. Atop these spherical assemblies would be many monopropellant thrusters. Each thruster should be directed away normal to the sphere-like manifold. These thrusters would be mounted on such a structure in order to provide multi-directional attitude control. They would be electronically controlled so the thrusters would fire when needed, in order to supply the proper moment. During the flight there may or may not be (depending on the MPS) different stages of the MPS mounted below the location of the bottom of the lander. This is the purpose of the assembly on top of the lander. It would be used during the flight until all lower engines are jettisoned. This would lead up to the descent phase. Here the top spherical thruster assembly could be jettisoned and the lower assembly could take over.
This system would have the benefit of being able to provide thrusts in a numerous amount of directions in order to maintain proper orientation. During descent, each thruster would also have a downward component. This would provide thrusts that aid the MPS in the ability to hover. This could lighten the load on the MPS, which in turn could allow it to have a smaller engine. Drawbacks to this system include complexity of design. It requires a good deal of plumbing, valves, and electronic controls to manipulate those valves. Also, during the stages that lead up to descent the top assembly would be used. This vertical components of this system would provide forces in the opposite direction of the velocity vector. That would create a greater load on the Main Propulsion System. It would have to overcome those extra forces.
The second design that will be under analysis involves mounting thrusters on the body of the lander and the main propulsion systems. They should be placed in couples as far away along an axis as possible so that the largest moment can be produced. Since the shape of the lander is octagonal coupled thrusters could be mounted on opposing faces of the body an equal horizontal distance through the center axis of each. Coupled thrusters could also be mounted on opposing faces an equal vertical distance from the axis that goes through the center of both. These two pairs of couples would prevent rotation about the y- and x-axes, respectively. To prevent rotation about the z-axis coupled thrusters could be mounted on the top and bottom of the structure an equal distance from the vertical axis, in opposite directions. This system can handle the encountered attitude control during the descent phase, but needs help in order to prevent/correct rotation during the prior phases of the flight. This is true because the extra engines that are staged throughout the flight create a different geometry. The necessary attitude control could be done by using half of the thrusters mentioned in the descent phase. A thruster could be mounted on the bottom face of any or all staged engines to create a coupled moment with the thruster at the top of the lander. The same concept could be used by placing a moment on the side of an engine that will be staged so that it creates a coupled moment with the thruster mounted on the top of the side faces of the lander. This method could be used to constrain all degrees of freedom for any number of stages during the flight.
A benefit of this design is its flexibility. It can be designed to adequately handle any number of stages designed by the MPS. The placement of the thrusters would be simple once the geometry is specified, and all rotational degrees of freedom could be constrained. A drawback could result from the number of thrusters that may be used, again , depending upon the number of stages. A separate fuel and plumbing system may be necessary for each thruster or for each coupled pair. If an MPS system is chosen whose staging does not affect the geometry, then this design would be quite adequate.
Now that a good conceptual design has been achieved the task at hand is to come up with more specific information. A specific launcher and the engine types must be selected for all thruster assembles. From here the size of the fuel tanks must be calculated. Coinciding with the size of the tank, is the amount of fuel required. If bi-propellant engines are chosen, plumbing systems for both the oxidizer and the propellant must be designed.
One question is what type of fuel should be used? Two types of fuels are solid and liquid. If a liquid rocket system is chosen, the propellants are stored in tanks and fed on demand into the combustion chamber by gas pressurization or a pump. More room needed for storage tanks must be considered in the design. If solid rockets are used, the fuel is typically powdered aluminum and the oxidizer is ammonium perchlorate. The solid rocket cannot be restarted, yet it is more reliable than liquid engines. Will solid fuel be better than liquid? Or is it a trade off for more mass being needed?
Another question to take into consideration is how much fuel will be needed for the mission? This is determined by what size engines, thrusters, storage tanks, etc. Determining the location of the thrusters will involve integration with the Structures Group.
The next step in the design process for the Propulsion Group is to take the conceptual designs and further develop them by answering the questions stated above and ones yet to be asked. After completing a more detailed study of each design, an optimal design will be chosen.
Wednesday, 31-Dec-1969 18:00:00 CST
CSR/TSGC Team Web