Selenium Technologies has been conducting preliminary design work on a manned lunar lander for use in NASA's First Lunar Outpost (FLO) program. The resulting lander is designed to carry a crew of four astronauts to a prepositioned habitat on the lunar surface, remain on the lunar surface for up to 45 days while the crew is living in the habitat, then return the crew to Earth via direct reentry and land recovery. Should the need arise, the crew can manually guide the lander to a safe lunar landing site, and live in the lander for up to ten days on the surface. Also, an abort to Earth is available during any segment of the mission.
The main propulsion system consists of a cluster of four modified Pratt and Whitney RL10 rocket engines that use liquid methane (LCH4) and liquid oxygen (LOX). Four engines are used to provide redundancy and a satisfactory engine out capability. Differences between the new propulsion system and the original system include slightly smaller engine size and lower thrust per engine, although specific impulse remains the same despite the smaller size. Concerns over nozzle ground clearance and engine reliability, as well as more information from Pratt and Whitney, brought about this change.
The power system consists of a combination of regenerative fuel cells and solar arrays. While the lander is in flight to or from the Moon, or during the lunar night, fuel cells provide all electrical power. During the lunar day, solar arrays are deployed to provide electrical power for the lander as well as electrolyzers, which separate some water back into hydrogen and oxygen for later use by the fuel cells. Total storage requirements for oxygen, hydrogen, and water are 61 kg, 551 kg, and 360 kg, respectively.
The lander is a stage-and-a-half design with descent propellant, cargo, and landing gear contained in the descent stage, and the main propulsion system, ascent propellant, and crew module contained in the ascent stage. The primary structure for both stages is a truss, to which all tanks and components are attached. The crew module is a conical shape similar to that of the Apollo Command Module, but significantly larger with a height and maximum diameter of 6 m.
1.0 Introduction1.0 Introduction
As mankind advances toward the permanent settlement of space, NASA finds it necessary to establish a lunar habitat capable of supporting life for extended periods of time. These extended duration missions may last anywhere from 14 to 45 days. Essential to the success of this habitat is a spacecraft capable of transporting a given set of crewmembers and cargo to and from the habitat. Selenium Technologies has been conducting preliminary design work on this lunar lander according to the requirements set by NASA in the Request for Proposal received in late January. The lander will provide the crew with the necessary transportation, life support, and cargo space necessary for each mission.
The system will rely on a Heavy Lift Launch Vehicle (HLLV) to reach low Earth orbit (LEO). Four crewmembers and limited cargo will be transferred to a lunar orbit. The spacecraft will be able to descend to any predetermined location on the surface of the Moon while providing the capability of redesignating the landing aim point during descent. The craft will provide the four crewmembers with life support for up to 10 days on the lunar surface while they prepare the lunar habitat for use. After as long as 45 days on the lunar surface, the craft will ascend from the lunar surface and return to Earth using a direct reentry and with a land recovery.
2.0 Orbits2.0 Orbits
The orbits subsystem provides the mission trajectory (i.e., Æv
burns required) which must be established before the sizing of
the other subsystems can be completed. When Æv burns are
known, the propulsion system can find the required fuel mass/volume
and the structures subsystem can begin sizing. This section of
the report will outline some important requirements and constraints
on the trajectory of the spacecraft and detail a mission trajectory.
2.2 Requirements and Constraints
There are several major requirements for the orbits subsystem. Time of flight between the Earth and the Moon must be no longer than four days. The lunar lander must be capable of landing at any site on the Moon. The crew module must make a land touchdown at the end of the mission. The lunar lander must be able to abort at any phase of the mission.
Several constraints and considerations also exist. Nominal lunar landings will occur near the local lunar sunrise and nominal lunar liftoffs will occur near the local lunar sunset. These two conditions would provide a maximum period of lighting on the lunar surface for the mission.
Two assumptions are made in this mission scenario. First, the
FLO is assumed to have already been placed on the Moon's surface.
Second, it is assumed that the HLLV will be capable of lifting
a 200 metric ton payload into a 185 km altitude orbit.
2.3 Mission Trajectory
The mission trajectory chosen for our lunar lander is the minimal-energy
trajectory detailed in NASA's FLO Conceptual Flight Profile document1.
The approximate Æv burns that are listed in the following
sections come from that document.
2.3.1 Earth-to-Moon Trajectory
Figure 2.1 shows the trajectory from the Earth to the Moon for
a mission to Mare Symthii. On December 5, 1999, a HLLV carrying
the lunar lander will launch from Kennedy Space Center. The HLLV
will boost the lunar lander and a lunar injection stage into a
185-km altitude parking orbit with a 33 degree inclination. At
the first injection opportunity, the lunar injection stage will
perform a 3140 m/s Æv burn to place the lunar lander on
its four-day transfer trajectory. For midcourse corrections along
the way, a Æv of 30 m/s is budgeted.
At the end of the outbound transfer trajectory, the lunar lander
will make an 830 m/s Æv burn to circularize around the Moon.
The altitude of the temporary parking orbit around the Moon will
be 100 km. When the appropriate phasing is reached, the lunar
lander will make a 20 m/s Æv burn to deorbit. During the
powered descent phase, the lunar lander will make a total of 1850
m/s Æv burn. Nominal touchdown will occur on December 9,
1999. The Æv burns for the major events on the outbound
trajectory are summarized in Table 2.1.
2.3.2 Moon-to-Earth Trajectory
Figure 2.2 shows the trajectory from the Moon to the Earth.
After a 42 day stay (45 days for contingencies) on the Moon, the
crew will liftoff from the Moon's surface in the ascent stage,
leaving the descent stage on the surface of the Moon. The powered
ascent phase of the mission will require a total of 1830 m/s Æv.
Once the ascent stage reaches an altitude of 100 km, it will
make a 20 m/s Æv burn to circularize the orbit.
When the phase conditions are met between the ascent stage and
the Earth, the ascent stage will make an 840 m/s Æv burn
to begin the transfer to the Earth. For midcourse corrections,
a Æv of 30 m/s is budgeted. As the ascent stage nears
the Earth, the crew module will separate from the rest of the
ascent structure. The crew module will make a direct re-entry
into the Earth's atmosphere. When atmospheric reentry begins,
the crew module will be traveling at a relative velocity of approximately
10.5 km/s. The crew module will follow an Apollo-type reentry
profile. After an approximately 15 minute reentry through the
Earth's atmosphere, the crew module will deploy parachutes and
fire retro-rockets (with approximately 20 m/s total Æv burn)
to make a soft land touchdown. Nominal touchdown will occur on
January 24, 2000. The Æv burns for the major propulsive
events in the return trajectory are summarized in Table 2.2.
|Lunar Orbit Circularization|
2.3.3 Æv Budget
The Æv burn numbers given in the previous two sections
correspond to the particular Earth-Moon geometry at the time of
launch. Table 2.3 shows the approximate Æv numbers corresponding
to the worst-case geometry between the Earth and the Moon.
2.4 Free Return Trajectories
On a lunar free-return trajectory, if the spacecraft is unable to make the Æv burn to circularize its orbit around the Moon, only a minimal amount of Æv is required to place the spacecraft on a return trajectory to the Earth. Out of concern for the criticality of engine failure in our preliminary single-engine lunar lander design, we studied free-return trajectories as an alternative to the minimal energy trajectory proposed in the FLO Conceptual Flight Profile document.
Several reasons prompted us to choose the minimal energy trajectory
over the free-return trajectory. The redesign of our lunar lander
from one engine to four engines decreased the criticality of single-engine
failure. Examination of typical Apollo Æv burns and preliminary
free-return trajectory analysis based on a computer algorithm
described in Battin2 showed an increase in
total mission Æv budget of 200-300 m/s over the total mission
Æv budget for the minimal energy trajectory, creating unacceptable
increases in mission mass estimates. NASA trade studies show
that any mass savings gained by the faster transfer time of the
lunar free-return trajectory are lost in the increase in propulsive
mass. In addition, the minimal energy trajectory has larger launch
and TLI windows than the free-return trajectory, allowing for
more flexibility in mission scheduling and execution3.
1. Langan, Michael P., et al., Mission Analysis Section. First
Lunar Outpost (FLO) Conceptual Flight Profile. Engineering
Directorate, Systems Engineering Division; NASA JSC, June 1992.
2. Battin, Richard H. An Introduction to the Mathematics
and Methods of Astrodynamics. AIAA Education Series: 1987,
3. Cockrell, Butch, NASA Project Manager -- SEI Lunar/Mars Flight Systems, telephone conversation. April 20, 1993.
3.0 Power Subsystem3.0 Power Subsystem
The power subsystem provides electrical power to the spacecraft
subsystems during all phases of the mission. The mission scenario
calls for a system that can operate at different power levels
and handle emergencies and aborts.
Input from the subsystems is required to form an accurate picture of the power requirements during the different phases of the mission. The power requirements available from the subsystems are used for a comparative analysis of different power options. During this analysis, it is assumed that the landing site does not have any nearby terrain features such as cliffs or mountains that would obscure the lander from sunlight during the lunar day. Appendix A shows the worst case breakdown of the mission, the power allotted to each subsystem, and the percentage of such power that is used during each mission phase. The total time for the mission is assumed to be 1291 hours. During this time the total amount of energy required is 2789 kilowatt-hours (kWh) at an average rate of 2.16 kilowatts (kW). Peak power (4.71 kW) occurs during the landing, takeoff and transfer burns, while minimum power (1.21 kW) occurs while the crew is in the FLO habitat.
Figure 3.1 summarizes the power requirements using a power requirements
timeline. The different shadings describe the different loading
conditions for the power system.
3.3 Criteria and Concerns
The preliminary criteria that were considered in the initial
selection of power system options were: mass, reliability, space
qualification, complexity, and safety. These criteria were only
used in an initial qualitative analysis of the power systems.
Once a preliminary selection of the power system options was
achieved, the mass of the system became the main driver for the
determination of the most adequate power system.
3.4 Power System Options
The initial group of power system options was qualitatively analyzed according to the criteria. Solar cells and rechargeable batteries, solar cells and regenerative fuel cells, and fuel cells were chosen for further analysis. A summary of the characteristics of the chosen systems is presented in this section, as well as a justification for the exclusion of other systems.
3.4.1 Solar Cells and Rechargeable Batteries
A schematic of a solar cell and rechargeable battery system is
shown in Figure 3.2. This system employs secondary batteries
to provide power during transit times from Earth to Moon, and
during lunar night. During the lunar day, solar panels are deployed,
taking over the power loads, and providing the energy required
to recharge the batteries.
3.4.2 Solar Cells and Regenerative Fuel Cells
Figure 3.3 shows a schematic of a typical solar cell and fuel
cell configuration. This system works similar to the previous
option in that fuel cells carry the load requirements during transits
and the lunar night. Solar cells carry the load during the day
and also provide power to electrolyzers for fuel regeneration.
The fuel cell-electrolyzer combination requires fuel, tanks,
piping, and control valves to control the passage of fluids throughout
the system. Hydrogen and oxygen are fed to the fuel cells, which
produce power and water. Some water is stored in a tank and electrolyzed
into hydrogen and oxygen for use during latter portions of the
3.4.3 Fuel Cells
The fuel cell system depends on the transformation of hydrogen
and oxygen for power production. The water that is produced during
the mission is given to the life support system. Excess water
is expelled from the lander. The nature of the system makes it
completely independent of the sun. Figure 3.4 shows a schematic
of the fuel cell system.
3.4.4 Other Options
The other preliminary options that were considered as possible power systems for the lunar lander included radioisotope generators (RTGs), solar dynamic converters, and large-scale nuclear reactors. These options were subjected to the initial criteria and were discarded.
RTGs use radioactive materials as heat sources. This heat is converted to electric power by means of thermoelectrics or a working fluid passing through a generator. The major drawback with RTG use is fuel availability. Current RTG designs use plutonium-238 (Pu-238) as the heat source and the entire stockpile of Pu-238 has been committed to interplanetary missions. The radioactive nature of RTGs could place the crew in a life-threatening situation should the power system malfunction, and would also draw large resistance from the political, public, and scientific communities.2
Solar dynamic converters use a solar collector to concentrate solar energy and heat in a fluid. The fluid is then placed through a dynamic cycle, producing electricity. Solar dynamic converters are a relatively new, unproven concept in space power generation; power levels are predicted to range from 10 to 40 kW. The main deterrent of this option is its relatively unproven technology and the complexity of its energy conversion system.3,4
Finally, large-scale nuclear power reactors were quickly discarded.
Their primary disadvantage is the inefficiency of the heat-to-electricity
conversion. Most space-tested large-scale nuclear power systems
have a thermodynamic efficiency of about five percent, resulting
in tremendous amounts of waste heat. Additionally, the crew would
require extensive shielding from the reactor's radioactive nature,
thus significantly increasing the mass of the system. Add to
this the political and public perception of nuclear power plants
and it is easy to see the difficulty in using a system like this.5
3.5 Final Selection Based on a Mass Analysis
The main driver for the final selection of an adequate power system is the mass of the system. The mission timeline and subsystem power requirements were used as the basis for this analysis. The first step was to determine the amount of energy required by the subsystems during each phase of the mission. The amount of energy required by the power system itself was then added to the energy requirements. Each power system was then sized using these requirements. The final results of this mass analysis are found in Table 3.2; the spreadsheets containing the specific analysis and intermediate results are found in Appendix A.
From Table 3.2, the choice of the solar cell and fuel cell power system as an adequate power system is evident. The solar cell and battery system is about five times more massive and the fuel cell system is double the mass. Although the solar cell and fuel cell system is complex and has not been space-qualified, the mass savings are too great to disregard.
After the selection of the solar cell and regenerative fuel cells,
a mission power timeline was developed. This timeline is different
from the previously presented power requirements timeline in that
it includes the power required for the power system itself during
the different mission phases. These requirements include power
for the pumps and valves in the fuel cell and electrolyzer sections
of the subsystem and also the solar cell power used for electrolysis.
The overall energy consumed during the mission now becomes 6342
kWh, used at an average rate of 4.91 kW. The maximum power output
of the subsystem occurs during the times when solar power is available
to perform the regeneration of the fuel, while the minimum power
still occurs when the crew is in the habitat. The final mission
timeline is found in Figure 3.5, while the mission breakdown is
found in Appendix A.
Further along during the development of the power system, it was decided that the fuel cell tanks were to be placed on the ascent truss, thus not permitting fuel cell power during the Earth reentry phase. The time allotted for reentry was small enough as to not affect the choice of the power option. However, a secondary battery system was chosen to provide power during this reentry phase. The specific batteries that were considered for this system are described in Appendix A. Four silver-zinc (Ag-Zn) rechargeable battery modules6 were chosen to provide power during the reentry phase. These have a relatively good shelf life and can be trickle charged to full capacity to replace any leakage. It is also important to notice that the fuel requirements were sized so as to provide enough energy for Earth return at any time during the mission.
The final subsystem sizing is found in Appendix A, a breakdown of the masses of the solar cell and regenerative fuel cell system is found in Table 3.3, and a final system schematic is found in Figure 3.6.
|Three Fuel Cells|
|Water (max. stored)|
|Pumps, Pipes, Struc.|
|Solar Arrays (GaAs)|
|Four Ag-Zn batteries|
1. McElroy, J.F., "SPE Water Electrolyzers in Support of Mission from Planet Earth," Journal of Power Sources, 1991, Vol. 36, No. 3, pp. 219-233.
2. "Spacecraft Power Generation," Spacecraft Subsystems (Student Spacecraft Subsystems Descriptions). Department of Aerospace Engineering and Engineering Mechanics, The University of Texas at Austin: January 1993, pp. 24-30.
3. Wertz, James and Wiley Larson eds., Space Mission Anaylsis and Design. Kluwer Academic Publishers and Microcosm, Inc.: Torrance, 1992, p. 393
4. "Spacecraft Power Generation," pp. 11-12
5. "Spacecraft Power Generation," pp. 31-34
6. Lunar Lander Design for the First Lunar Outpost, Engineering Directorate, Systems Engineering Division, NASA, JSC - 25896, p. 9-1.
4.0 Propulsion4.0 Propulsion
4.1 Introduction and Requirements
The propulsion subsystem performs all spacecraft propulsive activity.
Table 4.1 gives the activities, primary and secondary, which
are required in the mission scenario. The Æv estimates
given are from a worst-case Earth-Moon geometry trajectory analysis
performed by NASA.1
To perform all the requirements, the propulsion system must have certain capabilities, such as:
high thrust main engine(s) with throttle capability for landing
restart capability for main engine(s)
reliable reaction control system thrusters with proven multi-start capability and long lifetime
the RCS propulsion system must be able to start in zero gravity conditions
propellants which can be effectively stored for up to two months
In considering options for the propulsion system, the propulsion
system is subdivided into two parts: the main propulsion system
and the secondary propulsion system, or Reaction Control System
(RCS). The main propulsion system must perform all of the major
maneuvers and some abort maneuvers. The RCS must perform all
attitude control maneuvers and some abort maneuvers.
4.2.1 Main Propulsion System Options
Only propulsion systems which use liquid propellants were considered for this study, due to the need for restarting and throttling of engines. Hybrid solid propellant rockets with restart and throttle capability have been studied, but none have been flown into space, whereas numerous flight proven, restartable liquid propellant engines exist.2 Low thrust systems such as ion thrusters, arcjets and resistojets were also ruled out early since the landing and takeoff phases of the mission require a higher maximum thrust than these systems can provide.
Nuclear thermal rocket engines were ruled out primarily because of safety concerns. The required crew protection radiation shielding would add mass to the spacecraft structure. Also, current designs using nuclear thermal engines are most mass efficient for thrust ranges well above what is required for this project.3 Finally, given the current resistance from the public, scientific, and political communities to nuclear devices in space, any option requiring large quantities of radioactive material was ruled out for this project.
Monopropellant liquid propellant systems were not considered in depth because of their low performance compared to bipropellant options. For bipropellant systems, the two types studied were cryogenic and space storable. Cryogenic systems considered included liquid hydrogen/liquid oxygen(LH2/LOX) and liquid methane/liquid oxygen(LCH4/LOX). Storable propellant options included nitrogen tetroxide/monomethyl hydrazine(NTO/MMH). Other more exotic storable bipropellant combinations exist, but none have been extensively tested or flown.
The LH2/LOX option was ruled out because
of the requirement that the propellants be storable for up to
two months. The cryogenic characteristics of the liquid hydrogen
made propellant boiloff losses over two months significant. Considering
the two remaining options, the LCH4/LOX combination has significantly
higher performance than the NTO/MMH combination. Analyses showed
that this performance increase resulted in significantly more
mass delivered to the Moon's surface and to Earth reentry. Table
4.2 summarizes the results of some of the performance analyses
performed using the Æv estimates from Table 4.1.
Another factor considered which strongly influenced the design of the propulsion system was Mars Mission commonality. The project policy is to strive for commonality in systems in our project and the Mars Mission project. The Mars Mission is considering the use LCH4/LOX rocket engines as part of its propulsion system.
Based on the criteria discussed and the analyses performed, it
has been decided that the best option for this mission is a pump-fed
LCH4/LOX propulsion system. The decision matrix which was used
to come to this conclusion appears in Table 4.3. This system,
although much more complicated than the NTO/MMH system, has better
performance and Mars Mission commonality.
4.2.2 Reaction Control System Options
Due to the need for restart capability, solid propellant systems were not considered for the RCS. Also, only propulsion systems which rely on chemical decomposition or reaction were analyzed, because other systems, such as arcjet, resistojet, and ion types, have thrusts that are too low for this mission. Nuclear thermal options were not considered for the same reasons discussed in section 4.2.1.
Only pressure-fed bipropellant and monopropellant systems were deemed feasible for this mission. Pump-fed systems were not considered because the added complexity would have reduced reliability dramatically, due to the large number of start/stop cycles the RCS must perform. Cryogenic propellants were not considered because these are used mostly in pump-fed systems and because no small cryogenic rocket engines are currently available.
A desire to use proven technology eliminates all space storable
bipropellant combinations except for NTO/MMH. The monopropellant
options considered include hydrazine and peroxide. Both hydrazine
and peroxide have lower specific impulse ranges than NTO/MMH4,
although they are simpler and have lower mass than the bipropellant
system. All the propellants considered are toxic, with the NTO/MMH
being a hypergolic combination as well. The toxicity and hypergolic
character complicate ground handling, but this propellant combination
is so common that this complication is not viewed as a significant
obstacle. The hypergolic character, however, was viewed as advantageous,
since it negates the need for an ignition system, thereby enhancing
reliability. The specifi impulse performance of the NTO/MMH systems
is up to 100 seconds higher than either of the monopropellant
systems. Based on this, and the extensive flight experience of
NTO/MMH systems (including Apollo), a pressure fed, NTO/MMH propulsion
system was selected for the RCS of both the ascent and descent
4.3 Rocket Engines
Based on the propellant choice, a modified RL10 has been chosen for the main propulsion system. The modified RL10 has a maximum thrust of 69.4 kN and a mass of 272 kg, and has both gimbaling and throttling capability. The maximum thrust level of the modified RL10 is such that at least two are required for landing. For redundancy, four engines have been baselined. If an engine fails, its opposite can be shut down, and the remaining two engines throttled up. In this case, the thrust vector would still pass through the center of gravity of the spacecraft. This is the only engine that the project knows of which has been tested with the LCH4/LOX propellant combination. The RL10 has the advantage of a long heritage, with many successful space firings on Centaur upper stages. A schematic diagram of the RL10 appears in Figure 4.1.5,6 Miscellaneous engine performance parameters are outlined in Appendix B.
The choice for the RCS engines was based on availability, flight
experience, mass, and thrust. For the RCS system, the tentative
choice is the Marquardt R42 engine for the descent stage RCS,
and the Rockwell SE-8 ablative thruster for the ascent stage RCS.
Table 4.4 gives some of the characteristics of these engines.
The choice of the R42 for the descent stage is preliminary, since
the 890 N thrust is only an educated guess for what the RCS thrust
requirement needs to be. The SE-8 is the same thruster used on
the Apollo command module. The SE-8 was chosen primarily because
it uses an ablative cooling method. This cooling method is important,
since the thruster will be buried in the structure of the crew
module and thus cannot be cooled radiatively.
4.4 Propellant Feed System Design
The propellant feed system must provide propellant to the rocket
engines at a specified pressure and flow rate during all propulsive
maneuvers. The feed system includes the propellant tanks, pressurant
tanks, propellant lines, valves, filters, pressure regulators,
turbines, pumps, and ground support equipment hookups.
4.4.1 Main Engine Propellant Feed System
The main engine propellant feed system must provide LOX and LCH4 to the engines during all major propulsive maneuvers. The system must accommodate at least eight start/stop cycles, and must be able to store the propellants for up to two months at a time.
The main engine propellant feed system uses turbine-driven pumps to provide high pressure propellants to the combustion chamber. The turbine is driven by methane used in the regenerative cooling of the engine. The propellant tanks are pressurized to 50 psia using a helium tank and regulator system. Helium-operated valves are used to isolate the propellant tanks from the engines between firings, and Pyro valves are used to isolate the propellant from the rocket engines during ground operations and launch, after which the Pyro valves are fired to open the lines. Relief valves are manifolded to each tank to prevent tank rupture due to overpressure. A diagram of a Pyro valve appears in Figure 4.2.
The descent stage main engine propellant feed system also includes
a helium pressurization system. The four oxidizer tanks and the
four fuel tanks are linked manifolds. Each manifold is connected
to a relief valve to prevent tank rupture due to overpressure.
These valves are also used to vent boiloff of propellants. Schematics
for the ascent stage and descent stage main propulsion systems
appear in Figures 4.3 and 4.4. For clarity, the helium tank,
lines, and valves necessary to operate the engine valves are not
shown on the schematic.
4.4.2 RCS Propellant Feed System
The RCS propellant feed system must provide propellant to the RCS thrusters for all attitude control burns and for some abort maneuvers. The RCS must accommodate numerous start/stop cycles (100+) and must be able to start and function in both microgravity and high acceleration situations (i.e. the surface of the Moon), and must utilize storable propellants.
The RCS propellant feed system uses a pressure-fed design. High pressure gaseous helium tank(s) is (are) used to store the pressurant gas, which flows through a regulator into the propellant tanks. Since the engines must operate during various attitude maneuvers, including roll, yaw and pitch, the propellant tanks employ a positive expulsion system, thereby insuring that the propellant is always at the tank outlet. The tanks have burst discs/relief valves to prevent overpressure, and latch valves are used to isolate the thrusters from the tanks and the tanks from the pressurant tanks in between maneuvers. Pyro valves are used during ground handling and launch to isolate the
pressurant gas from the propellant tanks and the propellant from the thrusters. Pyro valves are also provided in some places to circumvent a stuck-closed latch valve.
The spacecraft requires two RCS propellant feed systems, one for the descent stage and one for the ascent stage. The feed system for the ascent stage provides dual redundancy in many places to minimize single point failures. The descent stage feed system does not have as much redundancy, since the ascent stage must be used for abort maneuvers anyway.
A schematic diagram of the RCS for both the ascent stage and
descent stage appears in Figures 4.5 and 4.6
4.5 Propellant Budget
Using the rocket equation9, the ÆV
estimates from NASA's FLO report, and the baseline propulsion
system design, an approximate propellant budget was produced.
Table 4.5 presents the budget; Appendix B shows the tk! Solver
model used to calculate this budget.
4.6 Future Work
For future work, the attitude control requirements must be analyzed to more precisely determine the amount of RCS propellant required.
1. Langan, Michael P., et al., Mission Analysis Section. First Lunar Outpost (FLO) Conceptual Flight Profile. Engineering Directorate, Systems Engineering Division; NASA JSC, June 1992, p. A-1.
2. Wertz, James and Wiley Larson eds., Space Mission Analysis and Design. Kluwer Academic Publishers and Microcosm, Inc.: Torrance, 1992, p. 23
3. "Propulsion Subsystem Report and Database," Spacecraft Subsystems (Student Spacecraft Subsystems Descriptions). Department of Aerospace Engineering and Engineering Mechanics, The University of Texas at Austin: January 1993, p. 23
4. Sutton, George Paul, Rocket Propulsion Elements. Wiley: New York, 1992, p. 567
5. Wertz & Larson, p. 673
6. Personal Communication with C.D, Limerick, Pratt & Whitney.
7. Wertz & Larson, p. 646
8. "Propulsion Subsystem Report and Database," p. 81
9. Huzel, Dieter K. and David H. Huang, Modern Engineering for Design of Liquid-Propellant Rocket Engines. Washington, D.C.: AIAA, 1992, p. 12
5.0 Guidance, Navigation, & Control5.0 Guidance, Navigation, & Control
The guidance, navigation and control (GN&C) subsystem provides
the spacecraft with the ability to determine the attitude and
position of the spacecraft, calculate a path from one attitude
and position to another, and control the spacecraft into that
new attitude and position. This section of the report describes
the preliminary GN&C system.
5.2 System Description
Inertial measurement units (IMUs) are used for attitude determination and, where applicable, acceleration measurements. For redundancy, the lunar lander has two IMUs, with the computer determining which IMU data to process. Four star trackers, stationed 90 degrees apart from one another along the circumference of the capsule, are used to realign the IMUs when they drift. Two radar altimeters are used to sense altitude above the lunar surface during powered ascent and descent.
The lunar lander uses 32 hypergolic thrusters for attitude control as well as translational control. More information about these thrusters is given in the propulsion section of this report.
The lunar lander uses autonomous and ground-based guidance. Autonomous guidance is used during time-critical events in the mission, such as the lower stages of powered descent and ascent, while ground-based guidance is used for other events, such as midcourse corrections. Powered descent is by automatic guidance, with provisions for manned guidance override.
GN&C software is processed by on-board computers. The lunar
lander has multiple, independent computers for redundancy.6.0
Environmental Control and Life Support System6.0 Environmental Control and Life Support System
The Environmental Control and Life Support System (ECLSS) provides an atmosphere of tolerable pressure, temperature, humidity, and composition, food, water management for both sustenance and hygiene,waste management, and fire prevention and suppression. With such an important function, the ECLSS system is found in many different forms with varying capabilities and penalties.
The ECLSS is made from a combination of many components; however, systems in use today are usually classified as one of three types depending on the amount of recycling used. The "open" (or "open loop") system stores everything needed to maintain a compatible environment, and after use all of the resulting waste is either stored on board or jettisoned from the spacecraft. The "partially closed" system is similar to the "closed" system. In this case, though, Electrochemical Depolarized Cells (EDC) filter out atmospheric carbon dioxide (CO2) and concentrate it for either removal from or storage aboard the spacecraft. The EDC produces electricity and heat as by-products of this CO2 reduction process. All atmospheric moisture and most hygienic and potable (drinking) water is either stored directly or produced from fuel cells. Multifiltration is usually used to reclaim some of the waste water for hygienic (i.e., non-consumption) use. The "closed" system attempts to continually recycle all of the atmosphere and water necessary for life support. Commonly used elements include an EDC combined with a Sabatier reactor which converts CO2 to potable water with methane and heat as waste products. Atmospheric oxygen is replenished by electrolyzing recovered wash water which also provides hydrogen necessary to run both the EDC and the Sabatier reactor. The remainder of the wash and waste water (including urine) is typically recovered by vapor compression distillation and used for non-consumptive purposes.
All of the systems briefly described above share two concepts.
First, they all rely on the storage or ejection of solid waste
produced by the crewmembers. Solid waste is stored in many ways,
two of the most common being vacuum or freeze drying. Second,
every system presented relies on the storage of foodstuffs rather
than on board production. When food generation is self-contained
inside the system, this is known as a "bio-regenerative"
system. However, a bio-regenerative is not feasible for such
a short-duration trip as this mission.
Life Support requirements are based upon two criteria, the number
of people that must be supported and the length of time they will
have to be supported. NASA guidelines have clearly established
the former criterion for this mission. The system must support
four crewmembers. The latter criterion is not so well defined.
The NASA requirements state that outside of a maximum lunar transit
time of four days one way, the crew must be able to remain in
the lander for a minimum of 48 hours on the lunar surface, and
total lunar surface time is to be 45 days (one lunar day-night-day
cycle). However, since the spacecraft is to be designed for an
extended stay several factors come into play. First, to provide
the crew with a habitation while preparing the FLO for occupation,
life support capabilities in the lander have been extended to
ten days. This leaves a maximum of 35 days in which the crewmembers
will reside in the FLO, leaving the lander unoccupied. During
this time, the lander will have to remain at some minimum state
of readiness in case of emergency. Therefore, it is estimated
that the ECLSS will have to run at approximately 25% capacity
during this time in order to maintain a habitable atmosphere and
allow for a rapid restart to full capacity if needed. Finally,
when calculating mass, volume, power, and waste heat generated
requirements, a 20% safety factor is added to ensure an adequate
amount of life support capability.
6.3 Comparison of Systems
Preliminary sizing figures for the three different types of ECLSS systems described were obtained from a tk! Solver program.1 The inputs were the criteria mentioned above, namely four crewmembers and a support capability of 38 days arrived at from the requirements discussed above. Both the program used and the input and output variables with their values are included in Appendix C. A note about the program worth mentioning is that one of the output variables is the mass of spares and consumables. Spares in this instance refers to both disposable parts like filters and redundant components in case of equipment failure. Consumables refers to food, oxygen, and other similar stores besides system hardware and spares.
The results of the sizing routine are clear. Figure 6.1 compares
the total mass of each ECLSS option. The open system is the most
massive with a total mass of 5544 kg, consisting mostly of consumables.
The open system is 4.1 times as massive as the partially closed,
and 8.6 times as massive as the closed. Of all the sizing results,
this is the most significant. Mass in spacecraft terms translates
directly to launch cost. In other words the open system is roughly
8.6 times more expensive to actually launch from the earth than
Figure 6.2 compares system volumes. Again, the need to store
large amounts of consumables (primarily oxygen and hydrogen for
air and water) with the open system results in a larger required
volume. Preliminary sizing revealed that the open system required
a volume of about 35.14 m3, or 4.5 times that
of the partially closed and 9.6 times that of the closed system.
These results are important from a structural standpoint, because
the more volume the system occupies, the greater the support structure
needed to accommodate it. The increase in support structure means
an increase in structural mass, which translates again into an
increase in cost and possibly structural complexity.
Figures 6.3 and 6.4 show the down side to systems that recycle
part or nearly all of their wastes. Figure 6.3 is the comparison
of system power requirements. This is calculated solely upon
the number of crewmembers in the spacecraft and the resulting
values represent peak power required (i.e., all systems being
run simultaneously at full capability). In this case, the open
system, being simpler in concept and requiring less hardware,
requires less power, about 0.78 kW, or 79% of the partially closed
and 44% of the closed system. These numbers are important from
the perspective that power systems are limited in the energy they
can produce. However, most power systems should have little problem
handling the 1.79 kW required by the closed system. Waste heat
generated, again dependent only on number of crewmembers, had
the same trend as power consumption. As Figure 6.4 shows, the
open system generated only 0.84 kW of heat at peak use. This
is only 65% of the partially closed system's requirement and 54%
of the closed system's requirement. Note that these numbers pertain
only to the heat generated by the ECLSS system hardware directly
and do not include the heat generated by computers and internal
controls including lighting, which add significant amounts of
heat that must be dissipated.
Based on these numbers, a conclusion as to the type of system best-suited for this mission can be easily reached. Mass is the most important factor in overall spacecraft design. Even with the capabilities of the assumed HLLV, every kilogram sent into orbit and to the Moon is costly. The mass penalty combined with the proven ECLSS technologies used in the closed system make it hard to justify using an open system with 8.6 times the mass of a closed one. The same reasoning leads to the conclusion that the partially closed system is likewise less suited for this mission than the closed. Therefore, Selenium Technologies recommends the use of a closed ECLSS like the one described in the introduction to provide the life support functions for the crew.
Furthermore, Selenium Technologies recommends the use of a reduced pressure atmosphere of 34.5 kPa total pressure and partial pressures of 6.61-7.86 kPa O2, 0.14 kPa CO2, and 26.5-27.75 kPa N2. The reason for this is to minimize or eliminatethe amount of pre-breathing time crewmembers will have to spend prior to leaving or entering the crew module. Other atmospheric conditions that have been established to maintain crew comfort are a temperature range of 18.3-26.7 _C, a dew point range of 4.4-15.6 _C, and a cabin ventilation range of 0.27-0.73 km/hr.
Two other recommendations are also made. First, the airlock
on the crew module must be equipped with a vacuum disposal system
to prevent lunar dust from EVA suits and equipment from contaminating
the crew module and fouling the air revitalization system. One
final recommendation is the inclusion of adequate fire detection
and suppression in the module. Although fire does not tend to
spread in the zero-g environment that will be encountered during
transit, there is a potential fire hazard (especially electrical)
during the lunar stay. Therefore, both smoke detectors and hand-held
CO2 fire extinguishers should be included
in the crew module.
1. "Environmental Control and Life Support System," Spacecraft Subsystems (Student Spacecraft Subsystems Descriptions). Department of Aerospace Engineering and Engineering Mechanics, The University of Texas at Austin: January 1993
7.0 Active Thermal Control7.0 Active Thermal Control
The active thermal control subsystem maintains all the
components of the spacecraft within their temperature limits by
either ridding them of excess heat or providing them with additional
heat. This process is divided into three components: the acquisition
component, the transport component, and the rejection component.1
The first two components of the system are common for
any environmentally controlled structure, but since the lunar
day's environment is so hostile, a special system is required
for the rejection phase on the Moon.
This active thermal control system is designed to rid the spacecraft
of a total of 7.5 kW of thermal energy. The ECLSS and power subsystems
produce 1.5 kW and 3.0 kW of excess heat, respectively. The excess
heat of the other subsystems used to size the radiators is based
on previous studies of similar spacecraft.
7.3 Acquisition and Transport
The acquisition (the first function of the thermal bus) and the transport (the second function of the thermal bus) will use a separate loop system. The first loop of this system is the acquisition component. It consists of a one-phase water loop, which acquires the heat from (or gives heat to) all the components of the spacecraft whose temperatures need to be controlled.2 The advantages of using water as the working fluid are that it is non-toxic, has high specific heat and extremely large heat of vaporization. Feasibility studies have shown that, even though the single-phase liquid system is more massive and requires more power than a two-phase system, it is more suitable for our system. Its simplicity and reliability have been proven numerous times in the past.
The second loop of the separate loop system is the transport
phase. It consists of a two-phase ammonia loop. The advantages
of using ammonia as the working fluid are that it requires less
pumping power, has a small total weight, and requires smaller
line sizes. This loop transports the thermal energy to the rejection
system. The ammonia acquires the heat from the water in the acquisition
system. Within the transport phase, the ammonia changes from
a liquid to a gas. It is carried in this form to the rejection
The most difficult task for the active thermal control system
is to reject the excess heat. Since the lunar surface phase is
a more hostile atmosphere than the transit phase, a different
system is required for heat rejection during the lunar surface
During the transit period to and from the Moon, inflatable composite radiators will be used to reject excess thermal energy. Six radiators of two different sizes to maximize mass and volume efficiency will be located on retractable booms. During transit, the booms will be extended to their maximum length so that the radiators will get a maximum exposure to space and, hence, approach their maximum efficiency. The total mass of these radiators and booms is 102.5 kg. Their dimensions, amounts of thermal energy rejected, and ttemperatures of operation are as follows:
2 radiators with 1.5 m diameter, operating at 294 degrees Kelvin, rejecting 4.4 kW
3 radiators with 1.25 m diameter, operating at 275 degrees Kelvin, rejecting 3.3 kW
1 radiator with 1.5 m diameter, operating at either temperature, for redundancy.
The arrangement of the radiators on the spacecraft is shown in
7.4.2 Lunar Surface
During the lunar surface period, one vertical radiator along
with a parabolic shading device will be deployed to the lunar
surface to reject excess thermal energy. This system is shown
in Figure 7.2. The system's total mass, including deploying equipment,
is 207.55 kg. This mass is greater than for an unshaded system,
but the system's efficiency is much higher than any unshaded system
because its operating temperature of 286_ K is much lower than
for the 360_ K operating temperature of the unshaded system.
The total area of the radiator is only 15 m2. Therefore,
the efficiency of the shaded vertical radiator is much better
than for the shaded horizontal system because it is able radiate
from both sides of the radiator and take advantage of every bit
of the system's area. The parabolic shading device blocks planetary
infrared and reflects and focuses incident solar radiation above
the vertical radiator located in the trough to insulate the radiator.
The trough and radiator are positioned on the lunar surface so
that the radiator is parallel to the plane of the Sun's path.
This system is capable of rejecting 7.5 kW of thermal energy.
The system will be stored in the cargo area during transit to
and from the Moon. However, if so desired the system can be left
on the lunar surface to allow for more cargo room on the return
Both rejection systems are compatible with the acquisition and
transport systems. The transfer from using the transit phase
rejection system and the lunar stay rejection system should require
little effort, even with the bulky EVA suits. After analysis
of many available systems, the previously outlined active thermal
control system was shown to be the most suitable system for this
mission. Therefore, Selenium Technologies recommends that it
be used on the Extended Duration Lunar Lander.
1. Guerra, Master's Thesis, "Active Thermal Control", University of Texas at Austin.
2. NASA, Engineering and Configurations of Space Stations and Platforms. Noyes Publications: Park Ridge, New Jersey, USA, 1985.
3. Ewert, Petete, Dzenitis, Advanced Environmental/Thermal Control and Life Support Systems, Active Thermal Control Systems for Space Exploration for a Lunar Base. Society of Automotive Engineers, Inc., 1990.
8.0 Communications Subsystem8.0 Communications Subsystem
The communications subsystem provides communication links between
all the components of the mission (i.e., spacecraft, earth).
The communications subsystem needs to accommodate high and low
data rates required for transmission of video, voice, science
and telemetry, and command signals between the spacecraft and
the ground stations on Earth. It is estimated that a high data
rate of approximately 10 megabits/sec for Earth-Moon links is
needed, mainly for transmission of compressed high-rate video
8.2 Communications Alternatives
A decision matrix used to rank the six alternatives for the communication
subsystem is shown in Table 8.1. Alternative A uses frequencies
in the S-band and a low-gain, wide-beam antenna for communications
service during all phases of the mission. Alternative B uses
the S-band for communication service while the spacecraft is in
LEO, during descent to and ascent from the Moon, and during Earth
reentry. During transfer to and from the Moon and during the
lunar stay, the X-band is used for the communication link. Alternative
C is similar to alternative B except that C uses frequencies in
the Ka-band for the Earth-Moon link. Alternative D is also similar
to alternative B except that D employs the Ka-band frequencies
instead of the X-band frequencies during transfer to and from
the Moon and during lunar stay. During transfer to and from the
Moon and during lunar stay, alternative E generates optical links
for communications service. It employs frequencies in the S-band
while the spacecraft is in LEO, during reentry, and during
TRANSFER TO/FROM MOON
|Reliability (5)||(5) 25||(5) 25||(3) 15||(2) 10||(1) 5||(5) 25|
|High data rates (5)||(2) 10||(4) 20||(5) 25||(5) 25||(5) 25||(4) 20|
|Continuous coverage (5)||(5) 25||(4) 20||(3) 15||(2) 10||(2) 10||(4) 20|
|Compatibility w/ ground station (4)||(5) 20||(5) 20||(3) 12||(3) 12||(1) 4||(5) 20|
|Low rain and cloud attenuation (5)||(4) 20||(4) 20||(3) 15||(3) 15||(1) 5||(4) 20|
|Mature technology (4)||(5) 20||(5) 20||(4) 16||(4) 16||(2) 8||(5) 20|
|Transponder power requirement (3)||(2) 6||(3) 9||(3) 9||(4) 12||(5) 15||(4) 12|
|Antenna (telescope) size and complexity (3)||(4) 12||(3) 9||(2) 6||(2) 6||(4) 12||(3) 9|
|System Mass (2)||(1) 2||(2) 4||(2) 4||(3) 6||(5) 10||(2) 4|
Total # of Points
descent to and ascent from the Moon. Finally, alternative F uses the X-band during all phases of the mission.
The alternatives were ranked based on the requirements listed in the Table's left column. The requirements include: subsystem reliability, provision of high data rates, continuous ground station coverage by the antenna beam, compatibility with currently used communication networks, low rain and cloud attenuation of the communication links, mature technological development, transponder power requirement, antenna size and complexity, and subsystem mass.
These requirements are satisfied best by alternative F. Since this alternative uses frequencies in the X-band, the antenna beam width needs to be narrow in order to support high data rates with low power supply. However, a narrow beam may not provide enough coverage for a real-time, continuous link. Nevertheless, this problem can be solved either by employing a waveguide lens antenna that produces a single beam with multiple lobes, or a reflector with an offset switched feed array. A reflector with an offset switched feed array generates multiple beams or a single beam that is hopped or scanned over the Earth's surface.2
By using frequencies in the X-band, the communications subsystem
avoids the overcrowded S-band for communications service. However,
if the X-band link fails during any phase of the mission, the
fail-soft design permits autonomous switching to an omnidirectional
antenna and the S-band in order to reestablish the continuous
link during the remaining part of the mission, or until the X-band
link can be recovered.3
Table 8.2 shows the frequency ranges that are used by the communications
subsystem. In the X-band, frequencies between 7.145 GHz and 7.190
GHz are used for the uplink transmission, and frequencies between
8.4 GHz and 8.5 GHz are used for the downlink transmission. The
corresponding downlink-to-uplink carrier frequency ratio is 880/749.
The S-band frequencies would be used only if the X-band link
fails during the mission.4 The above carrier
frequency ratios and frequency ranges are different from those
used by commercial broadcasting services.
* Frequencies used in case X-band link fails during mission
8.4 Communications Architecture
The above frequencies and carrier frequency ratios are compatible with those used by the Deep Space Network (DSN)4 and the Defense Satellite Communications System (DSCS)5. The DSN supports the communication link with the spacecraft during all phases of the mission as shown in Figure 8.1. If, in case of an emergency, the DSN is not available, the DSCS can be used to provide the link.
In case the mission requires that the spacecraft land on the
far side of the Moon or in case of an emergency descent on the
Moon, a lunar farside telecommunications relay satellite may be
needed to provide the link between the spacecraft and the ground
stations on Earth. This relay satellite would be placed in a
halo orbit at the far side of the Moon, as shown in Figure 8.1.
The relay satellite is a component of the complex communications
network planned for the lunar and Mars missions of the Human Exploration
8.5 Recommendations for Future Work
The preliminary design of the communications subsystem should allow for further modifications. For example, if higher data rates are needed for the Moon-Earth link, it may be necessary to use higher frequencies in the Ka-band.
Also, communication networks other than the DSCS should be considered for backup support of the communication link in case the DSN is temporarily not available.
Furthermore, optical systems should be considered for establishing
optical communication links during some parts of the mission.
In general, optical links provide higher data rates than microwave
links. Also, optical communication systems are usually lighter
and smaller, and require less power supply than their microwave
counterparts.7 However, optical links are
seriously attenuated by rain and clouds, which may result in discontinuities
in the communication link8. Furthermore,
the technological development of optical systems for space applications
is still in the early stage. More work needs to be done before
optical systems will be able to provide reliable communications
service during manned space missions. Nevertheless, this mission
should be considered further at least for experimental establishment
of innovative communication links, such as those generated by
1. "Report of the 90-Day Study on Human Exploration of the Moon and Mars," NASA News. National Aeronautics and Space Administration: Washington, D.C., November, 1989, p. 5-20.
2. Wertz, James and Wiley Larson eds., Space Mission Analysis and Design. Kluwer Academic Publishers and Microcosm, Inc.: Torrance, 1992, p. 542.
3. "Report of the 90-Day Study on Human Exploration of the Moon and Mars," p. 5-19.
4. Wertz and Larson, p. 516.
5. Fthenakis, E., Manual of Satellite Communications. McGraw Hill: New York City, New York, 1984, p. 332.
6. "Report of the 90-Day Study on Human Exploration of the Moon and Mars," p. 5-24.
7. Hauptman, R., "High Data Rate Atmospheric and Space Communications", Proceedings of SPIE-The International Society for Optical Engineers, Vol. 996. SPIE-The International Society for Optical Engineering: Bellingham, Washington, September, 1988, p. 4.
8. Wertz and Larson, p. 550.
9.0 Structures9.0 Structures
The structures subsystem provides a mechanical support for the other subsystems on the spacecraft. The structure of the spacecraft is formed by the primary structure, which carries the major loads, and the secondary structure, which provides support for different spacecraft components.
The structures subsystem group is responsible for all the areas
of the design that are related to the structure of the spacecraft.
These areas include ascent and descent stage configuration, crew
cabin configuration, propellant tanks design, materials selection,
and mass estimates.
9.2 Lunar Lander Configuration
The overall lunar lander configuration for the mission, shown in Figure 9.1, is formed by the ascent and the descent stages. The ascent stage is embedded in the descent stage, which is left on the Moon after the mission has been completed. The descent stage consists of the descent truss structure, propellant tanks for the descent, and the cargo. The ascent stage is formed by the RL10 engines, ascent truss, ascent propellant tanks, fuel cell tanks, and the crew module.
The configuration of the lunar lander is driven by the desire
to have two separate stages which share one propulsion system.
The overall dimensions of the lunar lander are driven by the
constraints on the payload area of the HLLV. The height of the
lunar lander is 16.2 m and the diameter is approximately 11 m.1
9.2 Descent Structure
The descent truss is composed of cylindrical aluminum members, with titanium end fittings for additional support. The structure holds all the fuel and oxidizer necessary for descent. This fuel will be linked to the ascent stage, connecting to the RL10 engines. The truss also houses the mission cargo and extra area for any life support or power equipment that cannot be contained within the crew module or in the ascent truss.
Figure 9.2 shows a top view of the descent stage with the legs
deployed. As shown in the figure, the descent stage has a platform
on top of it so that the astronauts can walk around the module
when they are on the surface of the Moon. Mounted on this platform
are two solar arrays that will be deployed during the lunar suface
stay to provide power. These solar arrays are shown retracted
in the figure, but they will be extended outwards to form a square
with 6.2 m sides. There is also an elevator mechanism shown in
this figure. The mechanism allows the astronauts to descend to
ground level and go back up, without using the ladder that is
mounted on one of the legs. The mechanism uses a small platform
to go up and down that fits within the descent truss structure.
The platform is lowered or raised with cables that are connected
to winches mounted on top of the descent stage. The whole mechanism
is powered electrically, but there are handles in the winches
as a mechanical backup system in the case of power failure.
Figure 9.3 gives a top view of the outer truss. The structure
is octagonal, with the fuel and oxidizer tanks arranged in a ring
surrounding the inner diameter of the truss. The inner diameter
of the truss is 6.3 m, which gives the ascent stage 0.6 m clearance.
The outer diameter of the structure is 10.8 m in order to fit
within the HLLV cargo constraints. Attached to the truss are
four retractable lander legs, which give the descent structure
a diameter of 20.8 m when the legs are fully extended.
Figure 9.4 shows a side view of the descent structure. The fuel
and oxidizer tanks are arranged in a ring around the inner diameter
of the truss. The oxidizer and fuel tanks are placed contiguously
at the top of the structure and occupy a height of 4.2 m. The
bottom 2.5 m of the truss is allocated for cargo and the deployable
radiators during lunar transit. The cargo bay employs a pulley
system which lowers all the cargo to the lunar surface.
Figure 9.4 also shows the width and length of each section on
the descent truss. Both of these dimensions are 2 m long, providing
enough space to house the large oxidizer tanks.
9.3 Ascent Stage
The ascent stage can be divided into two major components: the
ascent truss structure with the engines and the propellant tanks,
and the crew module.
9.3.1 Ascent Truss Structure
The ascent truss structure is formed by the ascent truss, propellant
tanks, fuel cell tanks, and the RL10 engines. As shown in Figure
9.5, the total height of the ascent truss structure is 10.2 m.
This distance is given by the fuel cell and propellant volume
requirements, the height of the RL10 engines, and the necessary
ground clearance for the exhaust nozzle. With this design, the
total height of the truss is 4.3 m, divided in two sections of
1.2 and 3.1 meters. The top section contains the fuel cell tanks,
while the bottom section houses the propellant tanks. Finally,
the RL10 engines have a total height of 4.4 m, which leaves a
nozzle ground clearance of approximately 1.5 m.
The top view of the ascent truss shown in Fig 9.4 has two different cross-sections. The cross section on top shows the fuel cell tanks, with two hydrogen tanks, one oxygen tank, and one water tank. The cross-section on the bottom shows the propellant tanks, with two big tanks for the oxidizer and two smaller tanks for the fuel.
The ascent truss structure is a square with 4 m sides, as shown
in Figure 9.4. This figure also shows that the maximum width
of the ascent structure is 5.02 m, which gives approximately 0.6
m of clearance between the ascent and the descent truss structures
at the closest point between the two. The members that form the
ascent truss have been designed to carry major loads experienced
during the mission, including bending, torsion, and compressive
loads that produce buckling. Finally, the materials used in the
ascent truss are aluminum for the truss members and titanium for
the fittings that connect these members.
9.3.2 Crew Module Configuration
The crew module for the baseline lunar lander, as shown in Figure
9.6, is a capsule similar in shape to the proven Apollo Command
Module. The crew capsule measures 6 m in height and 6 m in diameter
at the base. The interior walls of the module will be mainly
composed of an aluminum honeycomb material with additional aluminum
support beams. The exterior of the cone is covered by HTP-6 tiles,
an advanced form of the Space Shuttle's protective tiles. The
base of the module is covered by an ablator, which is the primary
thermal protection during the reentry into the Earth's atmosphere.
As shown in Figure 9.7, the thickness of the cone protection
is 7.5 cm, while the ablator has a thickness of 15 cm since the
base of the crew capsule will reenter the atmosphere first and
absorb most of the extreme heating that occurs during the reentry.
The crew cabin interior is divided into two decks, with the command
deck on top and the habitat deck on the bottom. The command deck
measures 2.2 m in height and 1.6 m in radius at the base. This
deck has a hatch used by the astronauts while on Earth to get
in and out of the crew cabin, and it can also be used as an escape
hatch in case there is an emergency landing in water. The habitat
deck measures 2.8 m in height and 2.9 m in radius at the base.
This deck is where the astronauts stay during lunar transit and
before transferring to the FLO habitat. This deck has an airlock
2.1 m high and 1.6 m in diameter, which eliminates the need for
depressurizing the entire crew module at the beginning and end
of each EVA.
9.4 Ascent-Descent Connections
Figure 9.5 is a schematic of the connections between the ascent
and the descent stage. As can be seen in the top view of the
figure, the total number of connections between the ascent and
the descent in each connecting area is 16. These connections
are simply aluminum rods that connect the descent stage with each
grid point in the ascent stage, except for the center grid point.
The side view of Figure 9.5 shows the two connecting areas between the ascent stage and the descent stage, one at the top and the other one at the bottom of the structures. Since there are two connecting areas between the ascent stage and the descent stage, the total number of connections is 32. These connections are all tilted, as shown in the figure, in order to decrease the bending loads transmitted by the connecting rods to the descent structure.
Finally, the figure also shows the fuel lines connecting the
propellant tanks in the descent stage with the RL10 engines.
These lines are placed at the bottom of the
structure to reduce the length of piping, and they are tilted
to provide gravity feed in case the pressure feed system fails.
One of the major tasks of the structures subsystem is to analyze the different materials suitable for space applications and select the ones that offer the best results. Some of the material properties that need to be considered in the selection of the materials are2:
strength to density ratio
stress corrosion resistance
fracture and fatigue resistance
electrical and magnetic properties
ease of manufacture.
In our design, the materials considered for the primary structure were aluminum, aluminum-lithium alloys, steel, titanium, intermetallic titanium alumides, magnesium, beryllium, and composites. The material that was finally chosen for the primary structure was aluminum due to its many advantages. Some of these qualities are: high stiffness to density ratio, high ductility, excellent workability, high corrosion, non-magnetism, moderate cost, and availability in numerous forms. The primary disadvantage of using aluminum is its low yield strength. Since aluminum lacks the strength to act as fittings between structural members, titanium was chosen for use in these areas.3
For the secondary structure, several materials were chosen according to their suitability for particular applications. The material selected for the inner wall of the propellant tanks was titanium, since the lunar lander uses cryogenic propellants and titanium exhibits good characteristics at low temperatures. For debris protection, we decided to use several layers of aluminized mylar insulation, which could also serve as thermal protection. Foam insulation and Schjeldahl coating, which has a low absorptivity to emittance ratio, also provide thermal protection.3
Aluminum was selected to provide radiation protection for the crew module. The 7.5 cm thick aluminum used in the primary structure provides the radiation protection so no extra material is required. The astronauts can be exposed to the amount of radiation this aluminum allows for up to six months with no ill effects.4
Finally, the materials chosen for reentry protection are AVCO-5026
ablator and HTP-6 tiles. The ablator material will be placed
at the base of the crew module since the module re-enters the
Earth atmosphere bottom first. The HTP-6 tiles will be placed
in the other areas of the crew module, which do not experience
the high temperatures of the base during reentry. These tiles
are a new generation of Shuttle tiles and the ablator material
is basically the same that was used in Apollo.5
9.5 Propellant Tanks
Figure 9.7 shows the sizes of the propellant tanks used for the ascent and the descent stages. As can be seen, the size of the tanks varies for the oxidizer and the fuel, and also for the ascent and the descent stages. The wall thickness of 8 cm for all the tanks is necessary to reduce boil-off.6
The spreadsheet used to size the tanks for this mission is found
in Appendix D. The propellant tanks are sized according to the
volume that is needed for each segment of the mission. A cylindrical
design with hemispherical caps was chosen because of its advantages
over a spherical design. The cylindrical design with hemispherical
caps minimizes the transfer of energy to the propellants by reducing
the overall area to volume ratio. Note that even though the area
to volume ratio for a single cylindrical tank is higher than that
of a spherical tank, the cylindrical design allows the number
of tanks necessary to be minimized, which translates to a lower
overall energy transfer.
9.6 Mass Estimate
The overall mass estimate for the mission is shown in Table 9.1. This overall mass estimate is based on the masses of the different subsystems, and also on the masses of Apollo and the FLO Mission done by NASA.7
Table 9.1 shows only the total masses of the mission. The detailed
mass breakdown can be seen in Appendix D, which contains the spreadsheet
used to calculate the mass of the lunar lander.
|TOTAL MASS||(Post TLI)|
1. Langan, Michael P., et al., Mission Analysis Section. First Lunar Outpost (FLO)
Conceptual Flight Profile. Engineering Directorate, Systems Engineering Division;
NASA JSC, June 1992. p. 1-4.
2. Wertz, James and Wiley Larson eds., Space Mission Anaylsis
and Design. Kluwer Academic Publishers and Microcosm, Inc.:
Torrance, 1992, p. 392.
3. "Spacecraft Structures," Spacecraft Subsystems (Student Spacecraft Subsystems Descriptions). Department of Aerospace Engineering and Engineering Mechanics, The University of Texas at Austin: January 1993.
4. Chilton, Schultis and Faw. Principles of Radiation Shielding. Prentice-Hall, Inc.:
Englewood Cliffs, 1984.
5. Langan, et al., p. 7-2.
6. Langan, et al., Sec. 8.
7. Langan, et al. Sec. 3.
10.0 Project Management and Cost10.0 Project Management and Cost
The management structure for the project is shown in Figure 10.1.
The management team consists of a Project Manager (David Garza),
a Chief Administrative Officer (Matt Carter), and a Chief Engineering
Officer (Tony Ng). All design work is divided among four primary
divisions: Orbital Mechanics/Guidance Navigation and Control,
Structures, Propulsion / Power, and Life Support / Active Thermal
A Division Manager leads each Division, and reports to the Chief
Administrative Officer (CAO) and the Chief Engineering Officer
(CEO). The CAO and CEO in turn report to the Project Manager.
A Division Manager may contact the Project Manager directly,
but most work should filter through the CAO or CEO so that all
management responsibilities are evenly dispersed. All design
decisions must ultimately meet with the approval of the Project
Manager. The responsibilities of the top management and the
Division Managers, along with all project tracking information,
are listed below.
10.1.1 Project Manager
The project manager oversees the entire project, and acts as
the primary contact with the contracting organization. Overall
program tracking and scheduling are handled by the Project Manager
and administrative duties are handled jointly with the CAO. In
the event of any major design obstacles, it is the duty of the
Project Manager to make the necessary decisions needed to keep
the project on track.
10.1.2 Chief Administrative Officer
The CAO handles the overall project management and shares all
administrative duties with the Project Manager. Some of the management
duties of the CAO include scheduling design meetings, maintaining
a project notebook, tracking project costs, and acting as a link
between the Division Managers and the Project Manager. During
the absence of the Project Manager, the CAO acts as the presiding
10.1.3 Chief Engineering Officer
The CEO has the overall responsibility of resolving any technical
dilemmas which may arise, including the integration of different
subsystems, transfer of necessary data between divisions, and
acquisition of technical data from outside sources. If the CEO
cannot resolve an important issue, it is reported to the Project
Manager and the issue is handled jointly. The CEO also supervises
the technical progress of each division and acts as a technical
consultant to each Division Manager.
10.1.4 Division Manager
The Division Managers have the responsibility of overseeing each
division and insuring the completion of the tasks assigned to
their divisions. Division Managers must also resolve any technical
issues involving their divisions, as well as schedule division
work assignments. If the issue remains unresolved, the CEO is
contacted and the problem is analyzed jointly.
10.1.5 Project Tracking
A Gantt chart for the project is shown is Figure 10.2. This
chart gives a sequential listing of the proposed project schedule.
10.1.6 Changes After the Preliminary Report
The only major change in the team organization and scheduling
after the Preliminary Report was the shifting of one member of
the Life Support/Communications Group to the Structures Group.
10.2 Project Cost
The cost considerations for this project include personnel,
computer, and supply costs. The cost analysis is drawn from twelve
weeks of work. Figure 10.3 shows the current cost analysis for
the personnel costs. These personnel costs are based upon the
salaries provided by the Request for Proposal. The straight line
in the graph depicts the estimated personnel cost which was initially
laid out in the proposal. As can be seen, Selenium Technologies
is well below this initial personnel cost estimate. This is primarily
due to an over-estimation of the personnel costs.
The project's computer costs were based upon the use of Macintosh
hardware and software. The hardware costs are based on rental
costs, while software costs were estimated as an initial lump
sum. These software costs were "paid" within the first
week and account for the large initial jump in cost. After the
first week, the computer costs began to level out. Since all
software costs were paid initially (which accounts for the over-budgeting),
the computer costs were very near the proposed costs by the end
of the project.
The supply costs cover all the materials necessary for presentations
and company communication. These materials include photocopies,
transparencies, model, poster, and miscellaneous materials. Figure
10.5 shows the actual supply costs versus the estimated supply
costs. Because of the added expense of the model and the poster,
Selenium Technologies' supply costs are slightly over budget.
The total project cost is shown below in Figure 10.6. Although
supply costs have slightly exceeded Selenium Technologies' expectations,
the low personnel costs have kept the total project costs under
budget. Since the personnel costs are the largest project costs,
they had the most effect on the overall project cost.
Preliminary Mission Timeline and Power Analysis
Mission timeline, power and energy requirements
|S/C operation||Time (hr)||Power (kW)||Energy (kWh)|
|in cap day||120.000||4.865||583.829|
|in FLO day||216.000||2.199||474.892|
|in FLO night||336.000||2.279||765.601|
|in FLO day||288.000||2.199||633.189|
|in cap day||48.000||4.865||233.532|
|in cap night||72.000||4.945||356.057|
|Fuel Cell Analysis|
|18.85 kg/kWh||2.24 kg/kWh||H2 density||O2 density|
|H2 (kg)||O2 (kg)||70.8||1141|
|0.311||2.617||percent of system||0.5|
|33.591||282.674||Mass of Array||0|
Regenerative Fuel Cells
|Fuel Cell Analysis (regenerative system)|
|18.85 kg/kWh||2.24 kg/kWh||Intermediate masses||Parameters for Electrolysis||H2 density||O2 density|
|H2 (kg)||O2 (kg)||H2||O2||Fuel cell n||0.550||70.8||1141|
|26.306||221.367||FC time (hr)||107.917||87.796||738.818|
|0.026||0.215||Elec time (hr)||336.000||Total||826.613|
|1.644||13.835||FC Power||5.172||Supporting mass|
|0.311||2.617||29.612||248.972||Electrolyzes||5.492||percent of system||0.5|
|33.591||282.674||Fuel Cell power||2.580||Mass of Array||174.068|
|26.323||221.511||Total fuel mass|
|Rechargeable Battery Analysis|
|Primary System||Intermed.||Require.||Time for|
|S/C Ops||Time (hr)||P (kW)||E (kWh)||kWh||Ah||hr|
|in cap day||120.000||4.865||583.829|
|in FLO day||216.000||2.199||474.892||1058.721||37.811||336.000|
|in FLO nig||336.000||2.279||765.601||765.601||27.343||336.000|
|in FLO day||288.000||2.199||633.189|
|in cap day||48.000||4.865||233.532||866.721||30.954||336.000|
|in cap nig||72.000||4.945||356.057|
|Options - Secondary Batteries|
|# of cells|
Reentry Battery Analysis
|Reentry Battery Options||Ag-Zn||NiCad||NiZn||AgCad|
Final System Sizing
|Fuel Cell Analysis (regenerative system)|
|18.850||2.240||Intermediate masses||Parameters for Electrolysis|
|H2 (kg)||O2 (kg)||H2||O2||Fuel cell n||0.550|
|1.164||9.795||1st regen - post land|
|24.846||209.103||FC time (hr)||107.833|
|1.552||13.060||Elec time (hr)||336.000|
|0.000||0.000||0.000||0.000||2nd regen - post FLO night|
|FC time (hr)||336.000|
|38.300||322.330||38.300||322.330||Elec time (hr)||336.000|
|electrolysis period||FC Power||2.427|
|0.146||1.227||Tot. Fuel mass||Water consumption|
|Add for return||78.794||663.114||741.908|
|at all times||12.107||101.893||114.000|
|Add for contingency||7.222||60.778||68.000|
|Assumed % of FC mass||1.000||unknown|
|Cabling and switches|
|% of total||0.500||552.000|
|NEED TO ADD SOLAR ARRAY||Preliminary|
|MASS TO THIS NUMBER =>||Total mass||1104.000|
|Solar cell array analysis (based on GaAs solar cells)|
|Space solar intensity (W/m2)||1358.000||Bus voltage||28.000|
|Angle between sun and cell normal (rad)||0.000||Array voltage||33.600|
|Solar cell efficiency||0.180|
|Degradation factors||1.000||Parameters||fuel cell|
|total degredation||0.000||Array||Specific W/kg||25.000|
|degredation coeff. (%/yr)||0.003||Weight (kg)||503.919|
|time of exposure to rad. (yr)||0.085||Area based kg/m2||3.300||264.000|
|time of exposure to rad (hr)||744.000||Area||80.000|
|thermal coefficient (%/degC)||-0.003||Assume||2.000||square panels|
|Maximum op. temp. (deg C)||130.000||Area/panel||fuel cell|
|Reference temp. (deg C)||28.000||40.000|
|Packing Factor||0.900||Side length||6.325|
|Maximum Power Required||4575.400||Mass Moments|
|Regeneration of Eclipse Power source||Perpen to face||11958.744|
|Perpen to axis||10344.458|
|fuel cell||Mass of support and|
|Power required||12597.974||assumed of|
|Solar Array Area (m2)||76.883||array mass||0.500|
|TOTAL SYSTEM MASS|
|Fuel Cell System||1777.90778|
|Solar Array System||396|
Mass Timeline of Fuels and Water
|Cum time||power||S/C||Energy||P. source||H2||O2|
|100.917||5.491||LOI burn||0.458||fuel cells||52.673||443.286|
|227.917||7.428||in cap day||891.390||panels+rege||55.382||466.088|
|443.917||4.922||in FLO day||1063.062||panels+rege||63.584||535.116|
|779.917||2.149||in FLO night||721.974||fuel cells||25.284||212.785|
|1067.917||6.481||in FLO day||1866.571||panels+rege||55.367||465.963|
|1115.917||8.988||in cap day||431.415||panels+rege||60.381||508.159|
|1187.917||4.655||in cap night||335.189||fuel cells||42.600||358.512|
|1194.500||5.491||TEI burn||0.458||fuel cells||40.878||344.021|
|H20 prod.||sum||H20 to ECLSS||sum||H20 elec.||H20 stor.|
tk! Solver Model for Propellant Mass Calculations
tk! Solver Variables Sheet
St Input Name Output Unit Comment
200 m0 mt mass after launch into LEO
mf1 104.78364 mt mass after TLI burn
95 m02 mt mass after TLI stage separation
L mf2 44.649348 mt mass after LOI and Lunar descent
m03 26.201848 mt mass after sep. from Hab and LDS
L mf3 12.499688 mt mass after Lunar ascent and TEI burn
mass at reentry
3230 tlidelv m/s Æ V for TLI
912 loidelv m/s Æ V for LOI
1873 descdel m/s Æ V for descent
1852 ascdelv m/s Æ V for ascent
878 teidelv m/s
Æ V for TEI
L Isp1 509.35304 s TLI Isp
L 376 Isp2 s LDS Isp
L 376 Isp3 s
9.81 g m/s^2 Acceleration due to grav. at Earth
mtlidry 9521.6364 kg dry mass of TLI stage
15500 mdescdr kg dry mass of LDS
0 mhab kg dry mass of habitat
dry mass of LAS
1.623 gMoon m/s^2 grav. accel. at Moon surface
L Tdescmi 27069.245 lbf minimum thrust of descent stage
Tascmin 9560.1384 lbf minimum thrust of ascent stage
1 T_to_Wt Thrust to Weight Rat. at Lunar surf.
155680 maxthru N
mp1 95216.364 kg Propellant mass for TLI
L mp2 50350.652 kg Prop. mass for LOI and DESC
L mp3 13702.159 kg
Prop. mass for ASC and TEI
L 6 MR1 mixture ratio for first stage eng.
L 46 rhofu1 kg/m^3 density of fuel in stage 1
L 1140 rhoox1 kg/m^3 density of ox in stage 1
mox1 81.614026 mt mass of ox in stage1
mfu1 13.602338 mt mass of fuel in stage1
L vox1 71.591251 m^3 volume of oxidizer in stage 1
L vfu1 295.70299 m^3
volume of fuel in stage 1
L 3.6 MR2 mixture ratio for second stage eng.
L 445 rhofu2 kg/m^3 density of fuel in stage 2
L 1140 rhoox2 kg/m^3 density of ox in stage 2
mox2 39609.902 kg mass of ox in stage2
mfu2 11002.751 kg mass of fuel in stage2
L vox2 34.745528 m^3 volume of oxidizer in stage 2
L vfu2 24.725282 m^3
volume of fuel in stage 2
L 3.6 MR3 mixture ratio for third stage eng.
L 445 rhofu3 kg/m^3 density of fuel in stage 3
L 1140 rhoox3 kg/m^3 density of ox in stage 3
mox3 13030.168 kg mass of ox in stage3
mfu3 3619.4912 kg mass of fuel in stage3
L vox3 11.429972 m^3 volume of oxidizer in stage 3
L vfu3 8.133688 m^3
volume of fuel in stage 3
262 mboilof kg boiloff from descent stage
2947.5 mboilof kg boiloff from ascent stage
m02_pos 74.189745 mt
tk! Solver Rules Sheet
"first leg: translunar injection
* call rocket(m0,Isp1,tlidelv;mf1) "final mass after TLI burn
* m02=mf1-mtlidry-mboiloff1 "mass after TLI separation
"prop. reqd. for Leg1
"second leg: lunar orbit insertion and descent
* call rocket(m02,Isp2,loidelv+descdelv;mf2) "mass after LOI and descent
* m03=mf2-mdescdry-mhab-mboiloff2 "mass after sep. of desc. stage
"prop. reqd. for Leg2
"third leg: ascent to orbit from lunar surface and TEI
* call rocket(m03,Isp3,ascdelv+teidelv;mf3) "mass after ascent and TEI
* mreentry=mf3-mascdry "mass after ascent stage sep.
"prop. reqd. for Leg3
"calculate minimum propellant volumes: stages 1, 2, 3.
* call vol(MR1,rhofu1,rhoox1,mp1;mox1,mfu1,vox1,vfu1)
* call vol(MR2,rhofu2,rhoox2,mp2+mboiloff1;mox2,mfu2,vox2,vfu2)
* call vol(MR3,rhofu3,rhoox3,mp3+mboiloff2;mox3,mfu3,vox3,vfu3)
"estimation of dry masses for stages
* mtlidry=0.10*(m0-mf1) "estimation of dry mass of TLI
C mdescdry=0.1*(m02-mf2) "estimation of dry mass of descent
"estimation of dry mass of ascent s
"estimation of minimum thrust for descent stage and ascent stage
* call rocket(m02,Isp2,loidelv;m02_postLOI)
"compute the # of engines
tk! Solver VOL Function
Comment: volume calculation
Argument Variables: MR,rhofu,rhoox,mp
Result Variables: mox,mfu,vox,vfu
tk! Solver ROCKET Function
Comment: rocket equation
Parameter Variables: g
Argument Variables: m0,isp,deltav
Result Variables: mf
tk! Solver UNITS Sheet
From To Multiply By Add Offset Comment
m/s km/s .001
kg lbm 2.205
ft/s m/s .3048
mt kg 1000
m/s^2 ft/s^2 3.28083989501
N lbf .224809024733
m^3 ft^3 35.31
Tank Sizing Spreadsheet
|Total Volume||Total Volume|
|Volume(Safety||Factor 1.05)||Volume(Safety||Factor 1.05)|
|Volume 1 Tank||(2 Tanks Total)||Volume 1 Tank||(4 Tanks Total)|
|Oxidizer Tank||Fuel Tank||Oxidizer Tank||Fuel Tank|
|Oxidizer Tank||Fuel Tank||Oxidizer Tank||Fuel Tank|
|Total Weight||Total Volume||Volume 1 Tank|
Lunar Lander Mass Breakdown
Descent Stage Mass Breakdown
|Conditioning & Wiring|
|Plumbing, Valves, etc.|
Ascent Stage Mass Breakdown
|Conditioning & Wiring||0|
|Displays and Controls||37|
|Active Thermal Cntrl||22.7|
|Fuel Cell Hydrogen||67|
|Fuel Cell Oxygen||550|
Command Module Mass Breakdown
|Pressure Vessel Structure||1272|
|Heat Shield Substructure||2104|
|Conditioning & Wiring||400|
|Displays and Controls||82.7|
|Spares & Consumables||293|
|Active Thermal Control||450|
|Suits and Hardware||245.6|
|Filters, Cartridges, Etc.|
|Other Life Support||35|