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Team 02D

IPT 2008 White Paper

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Alternate Concepts White Paper

IPT 02D

Systems Engineering
Team Leader
Systems Engineer
Guidance, Navigation and Control
Thermal Systems
Structures and Mechanisms
Power Systems
Payload and Communications
ESTACA and Southern University Liaison / Nick Case,UAH
Morris Morrel, UAH
Travis Morris, UAH
Greg Barnett, UAH
Thomas Stewart, UAH
Adam Garnick , UAH
Katherine Tyler, UAH
John Grose , UAH
Adam Fanning, UAH
Seth Farrington, UAH
Mobility Concepts
Modeling
Modeling / McArthur Whitmore, Southern University
Robert Danso, Southern University
Sample Return Vehicle
Modeling
Modeling / Julie Monszajin, ESTACA
Sebastien Bouvet, ESTACA

Submitted By:

February 21, 2008

Submitted To:

Dr. P.J. Benfieldand Dr. Matthew Turner

Department of Mechanical and Aerospace Engineering

The University of Alabama in Huntsville

Abstract

The purpose for Phase 2 was to develop two alternatives to the baseline design and use an engineering selection system to pick the better concept. Team LunaTech used a true systems engineering approach to begin outlining all options, determining concepts, and selecting the best concept. From this process, two alternatives, Lunar Prowler and Daedalus were created based upon the level 1 requirements and figures of merit. Team LunaTech used Pugh’s Selection Criteria to grade each concept and select the winning concept. Each concept had a unique structural design: Lunar Prowleremploys aland on wheels concept similar to the Mars Science Laboratory design; Daedalus uses a four-legged lander with a mobility concept attached. Each concept had a unique power design: Lunar Prowler usesa nuclear power supply to power all of its subsystems and allow it to explore the permanently dark areas needed to satisfy the scientific goals of the mission; Daedalusutilizes a solar cell and Lithium-Ion battery system to power all of the required subsystems. However, this limits its ability to land in permanently dark areas. Therefore, a unique strategy of launching lunar penetrators to the dark sites was created to overcome this limitation. The better concept was determined to be Daedalus. This concept was chosen because ofits efficiency,cost, the absence of a nuclear power source, and ability to meet the required launch date of September 30, 2012.A major design feature that led to the selection of Daedaluswas the power system. Although a nuclear power source would solve many of team’s design issues, early calculations and researchhas shown that it is not a feasible option. This is for several reasons, such as cost, weight, and schedule. The Daedalus design will allow for successful completion of the year-long mission, and could be capable of extending its life to achieve more scientific goals.

Technical Description

1.0 Overview of Phase 2

For Phase 2, the individual Integrated Product Teams (IPT)worked independently to produce two alternative configurations to a Baseline Design. The deliverables for Phase 2 are a White Paper and an Oral Presentation. The White Paper compares the Baseline Concept, in this case the Viking Mars Lander,with two alternative concepts. This White Paper summarizes the strategy for selecting alternative systems, the qualitative and quantitative information to evaluate each idea, and the logical rationale for selecting one concept over the other.

1.1Specification Summary

The primary requirements given by the customer are as follows: conduct geotechnical research at twenty unique lunar sites (15 permanently dark, 5 light) using a form of mobility; therobotic lander is responsible for collecting valuable scientific data and relaying it back to Earth using a direct link or the Lunar Reconnaissance Orbiter (LRO) relay system; the lunar lander is to be designed to survive a maximum 30 day cruise period, lunar landing, and a 1 year operational period. For the requirement of visiting twenty unique sites, the design must be flexible and able to operate a mobility platform in both light and permanently dark regions. For the data relay requirement, it is LunaTech’s goal to ensure the customer that all data will be received as quickly as possible. This will require a communication architecture that is robust and quick. To survive 1 year on the lunar surface, the design must be built with only the extreme conditions in mind. LunaTech must design a concept that will ensure the customer that at any location on the lunar surface the concept will survive. These major requirements will drive the design, and LunaTech will spend most of the emphasis on meeting these requirements.

1.2Team D Approach to Phase 2

The key challenges that the Concept Description Document (CDD) and the Baseline produced were how to survive the lunar night, how to survive a mission life of 1 year, how to survive concept of operations, and the choice of which design would produce the least mission risk and most reliability. Determining the concept with the right balance of risk and reliability is important for this project. The more risk, the more likely failure becomes, but with more risk one may be able to save weight, cut costs, etc. The amount of assumed risk is very important to the design of the system.

Team LunaTech followed the approach outlined in figure 1 to create a concept. First, the team identified all of the customer requirements and determined how they affected each subsystem. Next, the team divided into subsystem teams and researched solutions to satisfy the customer requirements. Then, the team converged as a whole and compiled all of the options for each subsystem to create a trade tree for two concepts. Since the initial trade tree was too large to incorporate into this document, a trade tree flow down chart was created as show in table 1. These concepts were then analyzed and graded using Pugh’s Selection Criteria matrix as shown in table 5. The team then compiled the best options from each of the 2 concepts to develop the final concept that will be used in the next phase.

1.3Trade Tree Reasoning

Once, the trade tree was created, a system-wide trade study took place. This was conducted in a period of one week and required all of the teams to work with all of the information they could find. Given more time, a more thorough trade universe would have been created, and a more thorough tree pruning process would have occurred. The removal and subsequent reasoning for each part of the trade tree will be documented here. It will also be supported with some calculations found in the appendix.

Power Subsystem

The power subsystems were unique for both concepts. The power subsystem had three main options: solar-battery power, solar-fuel cell power, and nuclear power. For the Lunar Prowler, a Radioisotope Thermoelectric Generator (RTG) was chosen over a Stirling Radioisotope Generator (SRG) because of its low Test Readiness Level (TRL) compared to the RTG. Solar cells were removed for the Lunar Prowler because of the need for Lunar Prowler to travel into permanently dark areas where solar cells would be ineffective. The fuel cell options were also removed because of the emission of water as by-product, and the needed mass required for a fuel cell system. For Daedalus, the nuclear power sources were removed because of schedule constraints, budget constraints, and mass requirements. The fuel cells were removed for the same reasons as stated above. Solar cells and batteries were chosen as the power supply for Daedalus. TheLithium-Sulfur Dioxide batteries were eliminated due to the voltage delay during initial power up. The voltage delay could last as long as seconds or minutes. The Lithium-Ion batteries were chosen because of their high power output to weight ratio, and availability for space applications. The current Li-Ion battery is based of the Venus Express Orbiter design.

Thermal Subsystem

The thermal subsystem had two main options: passive and semi-passive. The semi-passive system was chosen due to the belief that a passive system would not maintain the critical internal temperature required to ensure the integrity of internal components and instrumentation. Heaters must be used to maintain the narrow temperature window that is necessary for the batteries and other components. The semi-passive system that was removed was due to the following reasons: Thermostat controllers for the heaters are not as dependable as solid-state controllers and require more redundancy, and louvers were rejected due to the negative flight history of mechanical actuation

GN&C

For the altitude sensor, LiDAR was chosen over radar because of the LiDAR’s higher accuracy and added ability to provide the lander with hazard avoidance. For the vehicle state sensor, an IMU was chosen over an IRU similar to the Viking Mars design because the IRU has gyros only, no accelerometers. The IMU has gyros and accelerometers, giving the computer more information on the vehicles state. Multiple IMU’s will be used to provide redundancy in case of IMU failure or large drift. For positioning, video camera was ruled out because of possible time delays and a desire for a more precise, completely autonomous landing. Lunar GPS was not used because the option is currently unavailable for use. The use of radar to provide TERCOM was not used for scene matching because DSMAC technology with a camera provides higher accuracy than TERCOM.

Structures

The structure subsystem had two decisions paths: hard or soft landing and three or four legs. For the Lunar Prowler, the number of legs did not apply. A hard landing was chosen for both concepts to minimize dust disturbance caused by the propulsion system. Four legs was chosen over three because of the increased stability needed to land on larger inclines found near craters. This also provides more rigidity for the loading experienced during launch and landing.

Mobility

The options for the mobility branch of the trade tree were: land on wheels, land on wheels plus penetrators, lander plus a rover, lander plus a rover and penetrators, lander plus multiple rovers and a lander plus penetrators. The lander on wheels with penetrators was removed because the LOW platform was designed to explore all permanently dark regions and rendering the penetrators useless. The lander with a rover was removed based on the removal of nuclear energy from the available power source for the lander. This eliminated the use of a single rover on the lander. Lander plus multiple rovers was removed based on the belief that there should be ample amount of time to reach each site with a single rover. The timeframe for each sample site was estimated to be 1.2 weeks per site. The lander on wheels was chosen for the Lunar Prowler concept. The lander with penetrators and lander with penetrators and rover were viewed as plausible options and were left on the trade tree.


Figure 1 – Outline of the Design Methodology [02-D]

2.0 Description of Concepts

The two concepts chosen are identified as Lunar Prowler, and Daedalus. Lunar Prowler is a lander on wheels concept that has mobile capability. After landing, it will deploy the Single Site Science Box (SSSB) that will conduct all the single site science goals throughout the year. After deploying the SSSB, the lander will begin to navigate to each of the 20 sample sites via an electric motor drive train powered by two Radioisotope Thermal Generators (RTG) to conduct scientific analysis at each site. Daedalusdiffers in that it has no mobility itself. The Daedalus concept uses penetrators to conduct the scientific analysis at each of the 20 sites. The penetrators are based on the Deep Space 2 penetrators used by NASA and will be launched from a light gas gun designed to use the existing lander helium supply. The Daedalus will also have a rover that will transport a Sample Return Vehicle (SRV) away from the lander. The rover will also obtain and load the sample for the SRV.

Both concepts have the ability to complete the mission and satisfy all the customer requirements if designed correctly, but they both differ from each other significantly. Team LunaTech began evaluating each of the concepts and identified the major attributes of each. Differences in each subsystem for each concept were evaluated and researched in detail by individual team members in their specific concentration. Each individual then brought their results to the team as a whole. Once all subsystems for each concept were understood, the team proceeded to construct a Pugh’s Matrix to compare the concepts.

Table 1– Trade Tree Flow Down for [02D]

LETS Trade Tree Flow Down
Lander Concepts: Land on Wheels (LOW) and Non-LOW
Power / Thermal / GN&C
Solar w/Battery / All Passive / Computers
Li-Ion Battery / Semi-passive / Solid-State Controllers / Navigation
Li-SO2 Battery / Thermal Switches / Terrain Matching
Nuclear / Heat Pipes / Sensors
RTG / VRHUs / Altitiude
ASRG / Thermostat Controllers / LiDAR
Solar w/RFC / Radiators with Louver / Radar
with RFC / VehicleState
with H2RFC / IMU
IRU
Positioning
Camera DSMAC
Radar TERCOM
Lunar GPS
Video Camera
Controls
ACS
MPS
Surface Mobility
Communication / Structures / Mobility
DLS / Materials / LOW
Relay / Composites / LOW+Penetrators
Storage / Aluminum / Lander+Rover
Titanium / Lander+Rover+Penetrators
Landing / Lander+Penetrators
Hard Landing
Soft Landing
Legs
3 Legs
4 Legs

Table 2– Initial Mass Budget for [02D]

Given Mass Values from CDD
System / Mass / % of Total Mass
Total Landed Mass / 997.4 / kg / 100.00
Propulsion Dry Mass / 64.6 / kg / 6.48
Propellant Mass / 159.5 / kg / 15.99
Helium / 2 / kg / 0.20
Sample Return Vehicle / 250 / kg / 25.07
Lander Dry Mass / 932.8 / kg / 93.52
Initial Mass Budget Derived from Historical Percentages
Element of Weight Budget / Est.%of Lander Dry Mass / Est. Mass Based on Col. (1) (kg)
Payload / 30 / 279.8
Structures and Mechanisms / 20 / 186.6
Thermal / 4 / 37.3
Power / 17 / 158.6
Cabling / 7 / 65.3
GN&C / 6 / 56.0
Communication / 6 / 56.0
Margin (kg) / 10 / 93.3
Total / 100 / 932.8

Table 3– Initial Power Budget for [02D]

Given Mass and Power Values for Payload Units
Unit / Mass (kg) / Power (W)
Stereo Imaging System (SIS) / 3.5 / 9.5
Mast (SIS) / 0.8 / 6
Drill and Drill Deployment / 20 / 30
Arm / 13 / 43
Scoop / 0.7 / 0
Penetrator / 1 / 0
Mass Spectrometer / 19 / 75
XRD/XRF / 2 / 5
Total / 60 / 168.5
Initial Power Budget Based on Payload Power Requirements
System / Est. % of Lander Power / Est. Power Based on Col. (1) (W)
Total Budget / 100 / 700
Payload / 24 / 168
Structures and Mechanisms / 1 / 7
Thermal / 25 / 175
Power / 5 / 35
GN&C / 18 / 126
Communication / 17 / 119
Margin / 10 / 70
Total / 100 / 700

2.1Baseline Concept: “Viking Mars Lander” [02-BL]

Power System

The source of power for the Viking Mars Lander consisted of two Radioisotope Thermoelectric Generators (RTG). Each RTG weighed 15.4 kg with a minimum electrical power output of 35 Watts. The RTGs were supported by four Nickel Cadmium batteries. Each Nickel Cadmium battery had a capacity of 8 Ampere-hours. These batteries were used as energy storage for periods when the peak RTG output was surpassed and additional energy was needed to meet the power requirements.

Thermal System

The baseline design chosen for this project was the Viking Mars Lander. The Viking Mars Lander thermal control subsystem (TCS) employed the use of an overall semi-passive technique. Most of the components required to operate the lander during the landed mission were located in a single thermally controlled compartment. Heat from the Radioisotope Thermoelectric Generators (RTGs) was distributed to internal components via an equipment mounting plate fabricated from a single piece of high thermal conductivity aluminum and thermal switches. The thermal coatings selected for the Viking Mars Lander were selected based upon the need for a high value of infrared emmisivity, ε, and a balanced solar absorptivity.

Guidance, Navigation and Control (GN&C)

For this baseline evaluation, only the terminal descent sequence for the Viking Mars Lander will be evaluated, since the Lunar Exploration Transportation system (LETS) GN&C takes over at 5km above the lunar surface. The Viking GN&C system provided a completely autonomous landing sequence. Sensors consisted of an Inertial Reference Unit (IRU) that was used to measure body attitude rates with respect to the inertial frame, and a Terminal Descent & Landing Radar (TDLR) that was turned on at 5km altitude to measure surface relative velocity. A guidance computer was used to calculate and output the correct command signals to the controls. Complex algorithms and calculations were done within the computer and are beyond the scope of this paper. The controls consisted of three main terminal descent engines. All three main engines were differentially throttled and gimbaled to control the vehicles pitch, yaw, and relative velocity.

Structures and Mechanisms

The structural package for the Viking Mars Lander, shown in figure 3, consisted of a triangular frame with three strut landing legs. The landing legs were equipped with shock absorbers and honeycomb crush pads that allowed the lander to absorb the impact load of a hard landing.

Payload/Communications

The Viking Mars Lander’s payload parallels that which is planed for the LETS. The Viking Mars Lander landed on Mars with no mobile unit and was fixed to the surface. Scientific instrumentation for the Viking Mars Lander consisted of an X-Ray Fluorescent Spectrometer (XRFS) for determining chemical compounds, camera systems for observation of the horizon, and a Soil Sample Reconnaissance Apparatus (SSRA) with a boom, collector head, and shroud unit, capable of collecting a variety of material elements. Some of these experiments will carry over to the LETS mission.