NASA-CR-192033 /1 

Senior Deign Project

USRA/NASA/VPI&SU 

S 1/b 

Design of a Pressurized Lunar Rover 

S Final Report 

24 APRIL 1992 

S

SUBMITTED BY: 

MANOJ BHARDWAJ 

VATSAL BULSARA 

DAVID KOKAN 

SHAUN SHARIFF 

S ERIC SVARVERUD 

RICHARD WIRZ 

DEPARTMENT OF AEROSPACE AND OCEAN ENGINEERING

VIRGINIA POLYTECHNIC INSTITUTE AND STATE UNIVERSITY

ADVISORS: 

PROFESSOR ANTONY JAKUBOWSKI

DEPARTMENT OF AEROSPACE AND OCEAN ENGINEERING

VIRGINIA POLYTECHNIC INSTITUTE AND STATE UNIVERSITY 

DAVY A. HAYNESS

SPACE EXPLORATION INITIATIVE

NASA LANGLEY RESEARCH CENTER 

(-L-1111 033) JEIGN OF A 

L 'RESSURILFU LUNAR ROVER Final 

port (Virginia Polytechnic Inst. 

nd State Univ.) 110 p

COLOR PHOTOGRAPH 

N93-13O16 

Unc I as 

'73/37 0141625
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Design of a Pressurized Lunar Rover 24 April 1992 

CONTENTS 

i.LIST OF TABLES ................................................ 4

ii.LIST OF FIGURES ..............................................4 

iii.ABSTRACT ................................................... 6

1.INTRODUCTION ................................................ 7

1.1 Why a Pressurized Lunar Rover (PLR) ? .......................... 7

1.2 General Performance Requirements .............................8 

1.3 Lunar Environmental Design Considerations ...................... 8

1.3.1 Temperature ....................................... 9

.3.2 Gravity .......................................... 9 1

1.3.3 Radiation ..........................................9

1.3.4 Micrometeoroids .................................... 10

1.3.5 Trafficability ......................................10 

1 .3.6 Dust ............................................11 

2.CONFIGURATION SELECTION .................................... 12

2.1 Initial Configurations .......................................12 

2.2 HLLV and Shuttle C Evolution Effects .......................... 14

2.3 Intermediate Configurations .................................. 14

2.4 Final Configuration ........................................ 15

2.4.1 Selection ......................................... 15

2.4.2 Design Features .................................... 15

2.4.3 Internal Layout And Features .......................... 21

2.4.3.1 Driving Area ............................... 21

2.4.3.2 Living Area ................................ 27

2.4.3.3 EVA Area ................................. 28

2.4.3.4 Work Area ................................. 31

3.MATERIALS SELECTION ......................................... 32

3.1 Thermal Characteristics ..................................... 32

3.2 Solar Radiation ........................................... 32

3.3 Thermal Mismatch ......................................... 33

3.4 Micrometeoroids .......................................... 34

3.5 Honeycomb Core (NOMEX) .................................. 36

3.6 Flexible Material .......................................... 36

3.7 Graphite/Epoxy Selection .................................... 36

3.8 Material Selection For Vehicle Substructure ....................... 39

3.9 Conclusions .............................................. 39

4.MOBILITY SUBSYSTEM .......................................... 41

4.1 Configuration Selection ..................................... 41

4.2 Final Configuration ........................................ 44

4.2.1 Saddle ........................................... 44

4.2.2 "Pinned Wheel" Concept .............................. 44

4.2.3 Suspension System .................................. 50

4.2.4 Steering System .................................... 54

4.2.5 Wheel/Tire Assembly ................................ 58

4.2.5.1 Electric Motor ............................... 58

4.3 Performance ............................................. 62

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Design of a Pressurized Lunar Rover 24 April 1992 

5. POWER SUBSYSTEM ............................................. 66 

5.1 Power Required ........................................... 66 

5.2 Types of Power Systems ..................................... 66 

5.3 DIPS ................................................... 70 

5.4 Closed Brayton Cyc!e (CBC) .................................. 70 

5.5 Secondary Energy Storage .................................... 75 

5.6 Power Management and Distribution (PMAD) ..................... 76 

6. THERMAL CONTROL SUBSYSTEM ................................. 78 

6.1 Options ................................................. 78 

6.2 Internal Thermal Control .................................... 81 

6.2.1 Electrical Equipment ................................ 81 

6.2.2 Power System ...................................... 84 

6.2.3 Heat Pipes ........................................ 85 

6.2.4 Internal Air Temperature ............................. 87 

6.3 External PLR Thermal Control ................................ 89 

6.3.1 Multilayer Insulation ................................ 91 

6.3.2 White Epoxy Paint .................................. 91 

6.3.3 Windows ......................................... 91 

6.3.4 Solar Cells ........................................ 91 

7. LIFE SUPPORT SUBSYSTEM ....................................... 93 

8. GUIDANCE, NAVIGATION, AND CONTROL SUBSYSTEM ............... 97 

8.1 Guidance and Navigation .................................... 97 

8.2 Control ................................................. 99 

9. COMMUNICATIONS SUBSYSTEM .................................. 100 

9.1 The S-Band System ........................................ 100 

9.2 Voice Links .............................................. 101 

9.3 Telemetry Link ........................................... 102 

9.4 Command Link ........................................... 103 

9.5 Television Link ........................................... 104 

9.6 EVA Link ............................................... 105 

9.7 Antennas ................................................ 107 

10. REFERENCES ................................................. 108

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Design of a Pressurized Lunar Rover 24 April 1992 

L LIST OF TABLES 

Table 2.1 - Final Configuration Specifications ............................. 18 

Table 2.2 - Mass Statement ........................................... 18 

Table 2.3 - Mobility Performance Characteristics ........................... 20 

Table 2.4 - Subsystem Power Requirements ............................... 20 

Table 3.1 - Properties of T-650/35/Epoxy ................................ 37 

Table 3.2 - Pressurized Lunar Rover Shell ................................ 39 

Table 3.3 - Materials Mass Statement .................................... 40 

Table 6.1 - Temperature Tolerances ..................................... 81 

Table 6.2 - Thermal Control Subsystem Mass Statement ...................... 92 

Table 7.1 - Life Support Subsystem Design Loads .......................... 93 

Table 7.2 - Life Support Subsystem Characteristics .......................... 94 

Table 7.3 - Life Support Subsystem Parameters ............................ 95 

ii. LIST OF FIGURES

Figure 2.1 - Single Cylinder Configuration ............................... 12 

Figure 2.2 - Multi-segmented Spherical Configuration ....................... 13 

Figure 2.3 - Frame Options ........................................... 13 

Figure 2.4 - Dual Cylinder Configuration ................................ 14 

Figure 2.5(a) - Final Configuration (top view) ............................. 16 

Figure 2.5(b) - Final Configuration (bottom view) .......................... 17 

Figure 2.6 - Internal Layout .......................................... 22 

Figure 2.7 - Forward Cylinder, Starboard View ............................ 23 

Figure 2.8 - Forward Cylinder, Port View ................................ 23 

Figure 2.9 - Driving Area ............................................ 24 

Figure 2.10 - Horizontal Visibility ...................................... 25 

Figure 2.11 - Vertical Visibility ....................................... 26 

Figure 2.12 - Living Area ............................................ 27 

Figure 2.13 - EVA Area ............................................. 28 

Figure 2.14 - Rear Cylinder, Starboard View .............................. 29 

Figure 2.15 - Rear Cylinder, Port View .................................. 29 

Figure 2.16 - EVA Ingress and Egress ................................... 30 

Figure 2.17 - Work Area ............................................. 31 

Figure 3.1 - Thermal Mismatch ........................................ 33 

Figure 3.2 - Micrometeoroid Protection .................................. 35 

Figure 3.3 - Graphite/Epoxy Candidates ................................. 38 

Figure 3.4 - Pressure Hull Cross Section ................................. 40 

Figure 4.1 - Steering System Options .................................... 43 

Figure 4.2 - Vehicle Substructure, Top View .............................. 45 

Figure 4.3 - Joint Schematic .......................................... 46 

Figure 4.4 - Wheel Group ............................................ 47 

Figure 4.5 - Obstacle Extremes ........................................ 48 

Figure 4.6 - Terrain Performance ...................................... 49 

Figure 4.7 - Vehicle Substructure, Front View ............................. 51 

Figure 4.8 - Suspension Isometric ...................................... 52 

Figure 4.9 - Double Wishbone Suspension ................................ 53 

Figure 4.10 - Maximum Suspension Displacements .......................... 55

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Design of a Pressurized Lunar Rover 24 April 1992 

Figure 4.11 - Ackermann Steering System ................................ 56 

Figure 4.12 - Ideal (Neutral Steer) Turn .................................. 57 

Figure 4.13 - Wheel Assembly ......................................... 59 

Figure 4.14 - Motor System Diagram .................................... 60 

Figure 4.15 - Multiroller Planetary Drive System ........................... 61 

Figure 4.16 - Pull Per Wheel Versus Wheel Slip ............................ 65 

Figure 5.1 - Power Profile: Survey Mission ............................... 67 

Figure 5.2 - Power Profile: Transport Mission ............................. 68 

Figure 5.3 - Power Profile: Search and Rescue Mission ...................... 69 

Figure 5.4 - Functional Diagram of CBC ................................. 72 

Figure 5.5 - DIPS Power Module Heat Source Unit ......................... 73 

Figure 5.6 - Placement of Primary Power System ........................... 74 

Figure 5.7 - Power Management and Distribution Functional Diagram ........... 77 

Figure 6.1 - Thermal Control Functional Diagram .......................... 79 

Figure 6.2 - Thermal Control Options ................................... 80 

Figure 6.3 - Thermal Loop For Electrical Equipment ........................ 82 

Figure 6.4 - Radiator with Active Louvres ................................ 83 

Figure 6.5 - Control Block Diagram of the Radiator with Active Louvres ......... 84 

Figure 6.6 - Heat Pipe Schematic ....................................... 88 

Figure 7.1 - Life Support Subsystem Functional Diagram ..................... 96 

Figure 8.1 - Inertial Guidance System ................................... 98 

Figure 8.2 - Optical Tracker .......................................... 98 

Figure 9.1 - Voice Link ............................................. 102 

Figure 9.2 - Telemetry Link .......................................... 103 

Figure 9.3 - Command Link .......................................... 104 

Figure 9.4 - Television Link .......................................... 105 

Figure 9.5 - Picture Quality Levels ..................................... 105 

Figure 9.6 - Extravehicular Link ....................................... 106

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Design of a Pressurized Lunar Rover 24 April 1992 

iii. ABSTRACT 

--

A pressurized lunar rover is necessary for future long-term habitation of the moon. 

The rover must be able to safely perform many tasks, ranging from transportation and 

reconnaissance to exploration and rescue missions. 

Numerous designs were considered in an effort to maintain a low overall mass and 

good mobility characteristics. The configuration adopted consists of two cylindrical pressure 

hulls passively connected by a pressurized flexible passageway. The vehicle has an overall 

length of 11 meters and a total mass of seven metric tons. The rover is driven by eight 

independently powered two meter diameter wheels. The dual-cylinder concept allows a 

combination of articulated frame and double Ackermann steering for executing turns. In an 

emergency, the individual drive motors allow the option of skid steering as well. Two wheels 

are connected to either side of each cylinder through a pinned bar which allows constant 

ground contact. Together, these systems allow the rover to easily meet its mobility 

requirements. 

A dynamic isotope power system (DIPS), in conjunction with a closed Brayton cycle, 

supplies the rover with a continuous supply of 8.5 kW. The occupants are well protected 

from the DIPS system's radiation by a shield of tantalum. The large amount of heat produced 

by the DIPS and other rover systems is rejected by thermal radiators. The thermal radiators 

and solar collectors are located on the top of the rear cylinder. The solar collectors are used 

to recharge batteries for peak power periods. 

The rover's shell is made of graphite-epoxy coated with multi-layer insulation (MLI). 

The graphite-epoxy provides strength while the thermally resistant ML! gives protection 

from the lunar environment. An elastomer separates the two materials to compensate for the 

thermal mismatch. 

The communications system allows for communication with the lunar base, with an 

option for direct communication with earth via a lunar satellite link. The various links are 

combined into one signal broadcast in the S-band at 2.3 GHz. The rover is fitted with a 

parabolic reflector dish for S-band transmission, and an omnidirectional antenna for local 

EVA communication. 

The rover's guidance, navigation, and control subsystem consists of an inertial 

guidance system, an orbiting lunar satellite, and an obstacle avoidance system. In addition, 

the rover is equipped with a number of external fixtures including two telerobotic arms, 

lights, cameras, EVA storage, manlocks, a docking fixture, solar panels, thermal radiators, 

and a scientific airlock. 

In conclusion, this rover at a minimum meets all of the design requirements and 

clearly surpasses them in the areas of mobility and maneuverability. 

(

6
Design of a Pressurized Lunar Rover 24 April 1992 

1. INTRODUCTION 

1.1 Why a Pressurized Lunar Rover (PLR)? Man's return to the Moon will involve 

a permanent manned lunar base and outposts. These facilities will be supported by manned 

and unmanned rovers. Pressurized rovers will provide safe and comfortable transportation 

to areas in the vicinity of these outposts. Furthermore, they will perform extended missions 

far from their base. A pressurized rover operating on the Moon may also provide experience 

and technology for similar operations on Mars. This paper describes a design of a pressurized 

lunar rover. 

By the time that the PLR will be implemented, lunar settlements will have reached 

the lunar base stage. There will be a network of lunar outposts and other unmanned 

facilities. The PLR will transport personnel and supplies at the base, and between bases and 

outposts. It will also serve in search and rescue operations for other rovers, and will perform 

survey and exploration missions. 

Since the late 1950's, many papers have been written on unpressurized rovers. These 

papers outline many ideas especially in the areas of mobility and power. After the success 

of the Apollo program, pressurized, as well as unpressurized rovers, have been designed for 

future lunar missions. Many mobility schemes such as walkers and track-driven vehicles 

have been considered. The most common type of pressurized rover consists of a cylindrical 

pressure hull propelled by four to eight wheels. This general configuration was used as a 

basis for comparison of the various design options. 

Alternatives to the basic cylindrical design can introduce advantages. In light of the 

rugged terrain that the rover will encounter, the performance of the mobility system is 

critical. Due to the high cost of transporting mass to the Moon, it is also important to 

minimize weight. Many of the missions required of the PLR take place far from the base 

so the rover must be very reliable and safe.

7
Design of a Pressurized Lunar Rover 24 April 1992 

The design philosophy involved comparing different types of configurations that 

introduced advantages over the standard single cylinder. The importance of a low mass has 

already been mentioned. Mobility is critical since the primary functions of the rover are 

transportation and exploration of remote areas. The lunar surface is an extremely varied 

terrain, and the rover must be able to handle numerous conditions. Safety and reliability 

are also critical, since there will be no immediately available assistance for rovers on medium 

to long range missions. Simplicity is important for ease of repair and maintenance in the 

lunar environment. 

1.2 General Performance Requirements. The following is a list of the key design 

requirements for a PLR (Haynes): 

1. A nominal operational radius of 500 km (1,000 km range) per mission (lunar day). 

2. A nominal operational radius of 50 km (100 km range) for lunar night operations. 

3. Able to support a nominal crew of four. 

4. A nominal operational time of 14 days per mission. 

5. An airlock to allow EVA and which is compatible with the lunar surface habitat. 

6. An emergency one-time range of 2,000 km with a crew of two. 

7. Able to support a crew of six in an emergency with no range requirement (lunar 

safe-haven). 

8. Storage and consumable provisions to support two EVA suits for 28 hours of use 

each per mission. 

9. Direct communications capability with the Earth (audio, visual, and data). 

10.PLR shall be able to tow utility trailers with a mass of up to two metric tons. 

1.3 Lunar Environmental Design Considerations. A great deal of knowledge of the 

lunar environment resulted from the intensive studies and experiences of the Apollo program. 

8
Design of a Pressurized Lunar Rover 24 April 1992 

The following sections examine features important to the design of a vehicle for lunar 

operations. 

1.3.1 Temperature. Essentially the Moon is devoid of atmosphere. Unlike the Earth, 

radiation to and from the lunar surface is not hindered by an atmospheric blanket. This 

causes severe temperature ranges and gradients. Temperature extremes on the surface range 

from -233°C within shadowed polar craters to 123°C in the equatorial regions. Mean 

temperatures are about 107°C during the day and -153°C at night (Heiken et al). 

Temperature is a critical consideration for many subsystems. The thermal control system 

must be able to accommodate this wide range of conditions. Furthermore, it is important to 

avoid thermal mismatches between bonded materials. 

1.3.2 Gravity. The gravitational acceleration at the lunar equator is 1.62 m/sec2, 

about one-sixth that of earth. This is critical to the design of the mobility system since 

weight is a primary variable in determining vehicle performance. The human factors 

associated with this lower gravity are also important. Apollo 12 astronauts explained that the 

characteristic "loping" gait seen in films was the most natural way to move; heel-to-toe 

"Earth" walking was more difficult and energy consuming (Heiken et al). It is therefore 

critical that the vehicle's inner layout accommodate the astronauts as they deal with this 

environment. In lunar gravity, a 0.4 to 1.25 meter vertical hop is easier than using 

conventional stairs. This allows steps to be much larger. Also, seats that support the knees 

and buttocks are often used considering the posture people maintain. 

1.3.3 Radiation. The Moon is exposed to many types of ionizing radiation that can 

be biologically harmful. The three dominant types are the solar wind, solar flare associated 

particles (also called solar energetic particles or solar cosmic rays), and galactic cosmic rays. 

These forms of radiation consist mainly of protons and neutrons with some heavier nuclei. 

9
Design of a Pressurized Lunar Rover 24 April 1992 

Radiation levels vary with time and usually reflect solar activity. Obviously, the vehicle must 

protect its occupants from these harmful types of radiation. 

1.3.4 Micrometeoroids. A meteoroid is a naturally occurring solid body traveling 

through space smaller than a comet or asteroid. Micrometeoroids are those meteoroids with 

diameters less than approximately one millimeter. These evaporate upon entering Earth's 

atmosphere but are an important consideration for lunar activities. From size and frequency 

distributions, it can be estimated that a micrometeoroid of about one milligram mass should 

be expected to strike the rover yearly, with smaller objects being more frequent and larger 

ones rarer. Velocities have been calculated to range from 13 to 18 km/sec. It has been 

suggested that two to three millimeters of a tough composite material would provide effective 

shielding against damage by micrometeoroids in the milligram mass range. Another option 

is to use a buffer zone behind a less tough material. This would cause the micrometeoroid 

to disperse so that it cannot penetrate the hull. 

The rarer impacts of larger meteoroids pose a more significant hazard, especially to 

critical components. These components should be protected (by placing them where they 

cannot be struck) to ensure their safety. 

1.3.5 Trafficabllity. Trafficability is defined as the capacity of a soil to support a 

vehicle and to provide sufficient traction for movement. Many elaborate walking machines 

were developed prior to actual landings because it was assumed that trafficability would be 

poor. However, from the experience of both Apollo and Lunakod missions, it is now known 

that almost any vehicle with round wheels can perform satisfactorily on the lunar surface, 

provided that the ground contact pressure is no greater than 7-10 kPa. The energy consumed 

by wheeled vehicles can be estimated. Important parameters include the wheel load, wheel 

footprint area, chord length of wheel ground contact, width of wheel ground contact, and 

various previously measured soil characteristics.

10
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Design of a Pressurized Lunar Rover 24 April 1992 

1.3.6 Dust. The lunar regolith has grainy characteristics very similar to silty sand. 

Many of the particles are sharp and glassy (Heiken et al). Accumulation and adhesion to 

equipment was experienced on previous missions. "Sandblasting" can occur, especially in the 

presence of takeoffs and landings. Exposed or delicate equipment will have to be protected 

from the harmful effects of lunar dust. Also, dust removal equipment must be used after 

extra-vehicular activity.

S

11
Design of a Pressurized Lunar Rover 24 April 1992 

2. CONFIGURATION SELECTION 

2.1 Initial Configurations. The first step of the design process was to establish the 

basic configuration of the rover. The major subsystems and their relationships had to be 

agreed upon before in-depth work could begin. Preliminary studies showed that the pressure 

structure shape and mobility subsystem should be selected first because these elements have 

significant effects on other aspects of the design. This led to the selection of four initial 

configurations. 

One of the first designs considered was a slightly longer variation on the single 

cylinder often seen in literature. This design has a low center of gravity and a simple interior 

(Figure 2.1). A second design used tracks for enhanced mobility performance but was ruled 

out due to its high mass and power requirements. A single spherical pressure hull was also 

considered because of its optimum volume per unit mass. However, significant disadvantages 

included a high center of gravity and the utilization of available interior space. Therefore 

the fourth option of a multi-segmented spherical configuration was also considered (Figure 

2.2).

Figure 2.1 - Single Cylinder

12
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S 

S

S

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Design of a Pressurized Lunar Rover 24 April 1992 

Figure 2.2 - Multi-segmented Spherical Configuration 

S

Uniframe 

Multi-segmented 

Figure 2.3 - Frame Options 

The frame option was basically a choice between a uniframe configuration and a 

multi-segmented one (Figure 2.3). The uniframe, a standard one-piece frame, would be the 

simplest to construct and the most reliable. However, it needs more ground clearance, and 

therefore a higher center of gravity which decreases stability. Multiframe designs consist of 

a train of multiple segments, which allow pivoting at their connections. Increased 

maneuverability is achieved because the vehicle can conform to terrain conditions.

13
Design of a Pressurized Lunar Rover 24 April 1992 

2.2 HLLV and Shuttle C Evolution Effects. The cargo bay diameter of the launch 

vehicle affects the rover design. By completing most of the rover's construction on Earth, 

in-space assembly can be minimized. The projected dimensions of the National Launch 

System Derived Heavy Lift Launch Vehicles (HLLVs) and Shuttle C limit the size and shape 

of the pressure hull (Cook). For the lunar phase of space exploration, the HLLV's 

dimensions will be 7.62 meters (diameter) by 27.4 meters (length), and Shuttle C's dimensions 

will be 6.7 by 22 meters. Only one launch is required to bring the mass of the rover to the 

moon, using either launch system. 

2.3 Intermediate Configurations. As a result of preliminary analysis, three 

intermediate configurations were considered. To utilize the mass advantages of a spherical 

pressure hull, a multi-segmented dual sphere design was considered. The dual sphere 

configuration avoided the HLLV limitations of the single sphere. Utilizing the interior space 

of a spherical hull is difficult. Therefore, a simpler dual cylinder configuration was also 

considered (Figure 2.4). It maintained the mobility advantages of a segmented configuration 

without a significant mass increase over the dual sphere. 

Figure 2.4 - Dual Cylinder Configuration

14
Design of a Pressurized Lunar Rover 24 April 1992 

2.4 Final Configuration. 

2.4.1 Selection. Establishing the final configuration was reduced to a choice between 

three options: a dual sphere, a dual cylinder, and the standard single cylinder. Both of the 

segmented frame designs used a flexible pressurized passageway to passively connect the two 

pressure hulls. The flexible passageway increased the mass and complexity of the vehicle, 

but the mobility advantages were significant. The dual cylinder was chosen as the final 

configuration due to its increased mobility over the single cylinder and its simplicity 

compared to the dual spheres. It is a compromise between a uniframe and a many-segmented 

frame.

2.4.2 Design Features. The final dual cylinder configuration can be seen in Figure 

2.5. Its specifications are shown in Table 2.1. Table 2.2 is the mass statement, Table 2.3 

displays the vehicle's performance characteristics, and Table 2.4 lists the primary power 

consumers. 

S

S 

I

S

S

15 

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Design of a Pressurized Lunar Rover 24 April 1992 

L1⁄4) PH 

Figure 2.5(a) - Final Configuration (top view)

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Design of a Pressurized Lunar Rover 24 April 1992 

Figure 2.5(b) - Final Configuration (bottom view)

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Design of a Pressurized Lunar Rover 24 April 1992 

Table 2.1 - Final Configuration Specifications 

Primary dimensions: 

Total vehicle length 11 m 

Cylinder length 5 m 

Cylinder diameter 4 m 

Flexible passage length I m 

Wheel diameter 2 m 

Wheelbase 7.5 m

Primary power source: Dynamic Isotope Power System (closed Brayton cycle) 

Secondary power source: Sodium-sulfur batteries with photovoltaic array 

External fixtures: 

1 1.2 meter parabolic reflector antenna 

I omnidirectional VHF antenna 

6 exterior storage bins and tool boxes 

4 visibility and maintenance cameras with lights 

4 driving lights 

2 robotic arms 

1 scientific airlock 

2 manlocks 

1 towing hitch 

1 docking fixture 

2 EVA seats 

1 high resolution landscape camera 

19M 2 directable solar collector array 

2 7.5 m2 directable thermal radiators 

4 second surface mirror radiators 

Table 2.2 - Mass Statement 

SYSTEM 

Structure and Pressure Vessel 

Inner Graphite/Epoxy Shell 

NOMEX Core (Honeycomb) 

Outer Graphite/Epoxy Shell 

Elastomer Layer (for Thermal Mismatch) 

Aluminum Coating 

Flexible Connection 

Saddle 

KEYLAR 29 Micrometeoroid Shielding 

Thermal Control System 

Insulation

MASS (kg) 

553 

84 

559 

130 

378 

238 

652 

240 

130 

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Design of a Pressurized Lunar Rover 24 April 1992 

Radiator 90 

Pumps 20 

Heat Exchanger 50 

Piping 100 

Refrigerant 50 

Power system 

Engines 200 

Structure 60 

Heat Source Blocks 200 

Radiation & Housing 500 

Power Distribution 40 

Sodium Sulfur Batteries 16 

Solar Collectors 18 

Mobility 

Tires (8) 240 

Motors (9) 36 

Gearing Sets (9) 36 

Suspension Groups (4) 60 

Steering

Articulated Frame Joint & Connection 50 

Environmental Control/Life Support 

Atmospheric Control & Supply 80 

Atmospheric Revitalization 20 

Fire Detection & Suppression 10 

Water Recovery & Management 120 

Waste Management 20 

Extra-vehicular Activity & Support 20 

Manlocks 230 

Internal Facilities 

Communications 25 

Guidance and Navigation 10 

Galley 70 

Personal Hygiene 90 

Emergency Equipment 30 

Drive Stations 80 

Work Stations 40 

Sleep Quarters 60 

External Fixtures 

Visibility & Maintenance Cameras (2) 8 

High Resolution Landscape Camera 4 

Lights 4 

Antenna is 

Telerobotics 22 

Drill Coring 10 

Winch 5 

Hitch 2 

EVA Suits 450 

Crew 360 

Payload soo 

TOTAL MASS 7,015

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Design of a Pressurized Lunar Rover 24 April 1992 

Table 2.3 - Mobility Performance Characteristics 

Maximum Velocity 

Nominal Power Requirement 

Maximum Allowable Power 

Maximum Gross Pull 

Maximum Vehicle Acceleration 

Soil Compaction Resistance 

Normal Soil 

Soft Soil 

Hard Soil 

Minimum Turn Radius (in percent of vehicle length) 

Neutral Steer 

Oversteer 

Range 

Maximum Gradient 

Wheel Sinkage 

Normal Soil 

Soft Soil 

Hard Soil 

Climbable Step Height 

Crossable Crevice

(4 wheels), 14.7 km/hr

(8 wheels), 29.4 km/hr

(4 wheels) 1.36 kW/wheel

(8 wheels) 0.68 kW/wheel

1.5 kW/wheel

900 N/wheel

0.64 m/s2 

22 N/wheel 

81 N/wheel 

4.0 N/wheel 

78.3%

60%

2,000 km

26.5° 

1.03 cm 

2.86 cm 

0.21 cm 

0.53 M 

1.7 M 

Table 2.4 - Subsystem Power Requirements 

SYSTEM 

Drive Motors 

Normal Operations 

Maximum 

ECLSS 

Operation (computers, communication, etc.) 

Thermal Control 

Average 

Maximum 

Internal Requirements 

(includes lights, galley, hygiene, etc.) 

External Lighting 

Experiments 

Miscellaneous

POWER REQUIRED (W) 

5,452

12,000

160

600 

800 

1,356 

1,500 

1,000 

1,000 

1,000 

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Design of a Pressurized Lunar Rover 24 April 1992 

2.4.3 Internal Layout And Features. The interior layout of the rover is shown in 

Figures 2.6-2.17. The 4-man crew occupies the rover for up to two weeks while performing 

various missions including driving, EVA, sampling and data collection, and other tasks. The 

interior will provide a comfortable working environment with ample radiation and solar flare 

protection. 

The interior is separated into four areas: driving, living, working, and EVA. The 

driving and living areas are located in the front section of the rover while the EVA and 

work areas are in the rear. The Environmental Control & Life Support Systems (ECLSS) 

are located overhead to provide an ample amount of mass overhead for general radiation 

protection. The power systems and other utilities are primarily located under the floor of the 

rover.

2.4.3.1 Driving Area. The driving area, located in the front of the rover, has two 

driving stations. It is elevated to give the driver a good view of the landscape, aiding 

navigation (Figures 2.7, 2.8 and 2.9). The window placement minimizes the drivers' blind 

spots allowing a good view of the lunar surface (Figures 2.10 and 2.11). A transparent sheet 

is stretched between two spools across the windows. When lunar dust accumulates on the 

sheet, a clean section is rolled down over the window, replacing the dirty one. Overhead 

video monitors give the drivers additional visual assistance. The seats can recline to a 

sleeping position and are free to translate and rotate to allow access to the side mounted work 

stations. These stations are used for vehicle guidance, navigation, control, systems 

monitoring, data management, communications, and mission-specific tasks. 

The telerobotics in the front of the rover are controlled from the drive stations and 

are equipped with cameras for visual aid (Figure 2.5). They are able to cover a large area 

from one stopped position. The robotic arms can perform many tasks such as scooping, 

scraping, drilling, and digging into the lunar surface for samples. The samples are collected 

21
Design of a Pressurized Lunar Rover 24 April 1992 

Top view: 4 meters

EVA 

Computer Manlock sit 

Stowage station etc. 

Lounge chairsl \ ____ Dust-offl Stowaoe 

ouch) area 

F-T`11 Docking fixture 

@RGal'iey

_(Fold away) exercise area 

lley Bust- offr Docng 

area_ control, 

Scientific Sample Sink Personal Mission Manlock airlock taoojno hygiene experiments EVA 

and seat 

storage

S

Figure 2.6 - Internal Layout 

jE 

0
Design of a Pressurized Lunar Rover 24 April 1992 

Airlock 

Figures 2.7 Forward Cylinder, Starboard View

2 Meters 

Airlock 

Figure 2.8 - Forward Cylinder, Port View

23
Design of a Pressurized Lunar Rover 24 April 1992 

t 

METERS 

Fijure 2.9 - Driving Area

24
Design of a Pressurized Lunar Rover 24 April 1992 

z 

U

- 

C 

C

25
Design of a Pressurized Lunar Rover 24 April 1992 

Ii 

.-

26
Design of a Pressurized Lunar Rover 24 April 1992 

by a scientific airlock which is also accessed by the drill corer located underneath the rover. 

After the samples are collected they are brought to the sample storage bin by a hand-powered 

conveyor belt. Here they are tagged and stored in a vacuum or in nitrogen-filled containers 

for future examination. 

S

2.4.3.2 Living Area. Immediately aft of the driving area is the living area (Figures 

2.7, 2.8, and 2.12). This allows the front of the rover to be an open area for crew meetings. 

S The lounge chairs can be assembled for use as beds or a couch. A sound/light curtain can be 

drawn around the lounge chairs for privacy. The compact galley and a multi-purpose sink 

are located next to the personal hygiene facilities. The personal hygiene room is also a 

shower, which is provided because of the possibility of several EVA excursions during 

missions. 

S

S

LIVING cIIID 

AREA

BIDS'LOUNGF (IIAIRS 

0

UFILITY S

2 Meters

Figure 2.12 - Living Area 

S

27
Design of a Pressurized Lunar Rover 24 April 1992 

2.4.3.3 EVA Area. The EVA area splits the work area so the exterior portion to the 

manlocks does not interfere with the vertical translation of the wheels (Figures 2.13-2.15). 

A manlock is a type of form-fitting airlock. By using a pressurized volume of 1.57 m3, the 

manlock conserves energy for EVA ingress and egress. A folding bridge and ladder allow the 

lunarnauts to avoid the wheel group (Figure 2.16). The Suits are stored inside the manlocks. 

A post-EVA dust-off area is provided to prevent lunar dust from contaminating the interior. 

a.

2 Meters

Figure 2.13 - EVA Area

28
Design of a Pressurized Lunar Rover 24 April 1992 

Figures 2.14 Rear Cylinder, Starboard View 2 Meters 

Figure 2.15 . Rear Cylinder, Port View

29
Design of a Pressurized Lunar Rover 24 April 1992 

Figure 2.16 - EVA Ingress and Egress 

A protection area is required to shield the occupants during periods of high solar 

flare activity. One study of space radiation protection examined the lunar regolith itself as 

a possible radiation shield for lunar base surface systems (Nealy et a!). The proposed design 

called for about a 0.5 meter thick shield, an unrealistic mass for a transport vehicle. More 

appropriate materials such as water and aluminum have been investigated. For galactic 

cosmic rays, 3 g/cm2 of water or 6 g/cm2 of aluminum would provide sufficient protection. 

For solar flares, 15 g/cm2 of water or 18 g/cm2 of aluminum are necessary (Townsend et al). 

However, the mass of a material dedicated solely for the purpose of providing radiation 

protection becomes unreasonably large for a crew of three or four. The solution is to use 

consumables such as water, radiators, and trash to provide shielding. Shielding on the order 

of 20 g/cm2 can be assumed sufficient for solar flare protection (NASA Memorandum 4170, 

Volume III). The EVA area is surrounded by relatively massive equipment such as the solar 

30
S

S

S

S

Design of a Pressurized Lunar Rover 24 April 1992 

collectors and thermal radiators overhead and the manlocks on either side. Therefore, the 

EVA area is used to protect the occupants during solar flares. 

A docking fixture to mate with the base is located in the very back of the rover. The 

base uses its airlock to transfer the rover passengers. The docking fixture allows direct 

S

passage between the rover and the base. 

2.4.3.4 Work Area. The work area is made up of several parts (Figures 2.14, 2.15, 

and 2.17). Computer and mission experiments stations are used to perform mission specific 

tasks. At the mission experiments station, samples and observations obtained during the 

mission can be analyzed. In-depth system and maintenance checks as well as data relays can 

be carried out at the computer station. Minimal controls which are located in the very rear 

of the rover are used to aid docking. Interior storage and emergency facilities are also found 

in the rear. A fold-away exercise area that can aid in charging the batteries is provided and 

can be easily stowed, depending on mission needs. 

2 Meters 

S Figure 2.17 - Work Area

31 

0
Design of a Pressurized Lunar Rover 24 April 1992 

3. MATERIALS SELECTION 

When selecting materials for the rover structure, the harsh lunar environment must 

be considered. Important criteria include large temperature extremes (-233°C to 1230C), 

thermal cycling with sharp temperature gradients, solar radiation, and micrometeoroids (see 

§1.0). Materials must be chosen that can withstand these harsh conditions and minimize 

mass. Also, the internal and external loadings on the rover will require the selection of 

extremely durable materials. High modulus, graphite/reinforced epoxy type materials are the 

leading candidates for most large spacecraft systems (Blankenship et al). The use of a 

graphite/epoxy composite has many advantages. The material is strong, lightweight, 

thermally stable, and like most other composites, can be tailored to meet particular needs. 

3.1 Thermal Characteristics. With large, divergent temperature extremes, the 

material on the exterior of the pressure hull will expand and contract. The magnitude of 

expansion is related to the material's coefficient of thermal expansion. Ideally, a material 

that has a negligible coefficient such as a graphite/epoxy should be used. 

Thermal cycling is a greater problem than temperature extremes. Repeated cycling 

over a prolonged period can cause severe microdamage, altering both the stability and 

stiffness of the graphite/epoxy (Blankenship et al). The particular microdamage that 

degrades the integrity of the material is called microcracking. A solution to this problem is 

to place multilayer insulation (ML!) over the graphite/epoxy. It minimizes the temperature 

cycling range by increasing the heat capacity of the shell. 

3.2 Solar Radiation. The ML! also provides an excellent defense against solar 

radiation. If the graphite/epoxy is exposed to radiation, chemical reactions begin to occur. 

These reactions severely alter the graphite/epoxy's material properties. Some of the adverse 

results include a reduction in the transverse strength and modulus of elasticity, increased 

32
Design of a Pressurized Lunar Rover 24 April 1992 

volatility, and a reduction in thermal stability (Sykes et al). By simply protecting the 

graphite/epoxy with ML!, the harmful effects due to solar radiation are reduced. 

3.3 Thermal Mismatch. There is a thermal mismatch between the ML! and the 

graphite/epoxy. Since the coefficients of thermal expansion are considerably different, their 

expansions are unequal. If the mismatch problem is not corrected, the ML! may crack or 

separate from the graphite/epoxy. The solution to this problem is to put an elastomer layer 

between the aluminum and the graphite/epoxy (Kander). The elastomer, having rubber like 

characteristics, will yield to the thermal mismatch (Figure 3.1). A good elastomer candidate 

is isobutylene isoprene because of its high continuous use temperature of 150°C. 

Note: Figure not drawn to scale. 

llj

Figure 3.1 - Thermal Mismatch
Design of a Pressurized Lunar Rover 24 April 1992 

3.4 Micrometeoroids. Ballistic impacts on the rover by micrometeoroids could prove 

catastrophic. A bumper shield system will be used to combat impact damage. The system 

entails the use of a bumper shield, spacing, and a back wall. The principle behind this 

scheme is that as a micrometeoroid breaks through the bumper shield, it disintegrates. The 

spacing behind the bumper provides room for the disintegrated particles to disperse. The 

back wall stops the dispersed material. The critical parameters to this design are the bumper 

shield thickness (ta), the spacing (S), and the back wall thickness (tb) 

According to NASA standards, the bumper thickness is 10 to 20% of the impacting 

particle's length (Wertz). For the lunar environment, the average micrometeoroid will have 

a diameter of one millimeter. Based on safety and manageability, the bumper shield will be 

made of the conventional Al 7075-T6 with a one millimeter thickness. 

Although the MLI and elastomer layers of the shell are the outermost, they do not 

serve well as the back wall. Therefore, the outer graphite/epoxy shell of 0.25 cm thickness 

will essentially serve as the back wall. 

According to the equation

tb = (Cmv)/2S 

where 

tb = back wall thickness = 0.25 cm 

m = micrometeoroid mass = 0.001 g 

C = 41.5 ± 14.0 

v = micrometeoroid velocity = 13 to 18 km/s 

S = spacing thickness (cm) 

the spacing thickness may be calculated (Wertz). Using a factor of safety of two, the spacing 

thickness was found to equal four centimeters. 

Due to the relatively high frequency of' micrometeoroid impacts in the lunar 

environment, rather than having one large bumper shield, a panelling scheme will be used. 

34
cm Spacing 

_T 

Design of a Pressurized Lunar Rover 24 April 1992 

This allows for individual panels to be replaced if considerable damage has occurred. Strips 

of KEYLAR 29 1-beams will be wrapped circumferentially around the rover. The aluminum 

panels will be fastened onto the 1-beams thus providing the four centimeter spacing (Figure 

3.2).

Al 7075-T6 Panel 

I II I

Ii

I I

Ii 

I I 

I

I I 

I

KEVLAR 

29 

I 

Ii II

"I-beam" I

I

Al 7075-T6 

bumper 1 mm

•L V Li1% i..' LIJa1II 

Figure 3.2 - Micrometeoroid Protection

35 
Design of a Pressurized Lunar Rover 24 April 1992 

3.5 Honeycomb Core (NOMEX). Since the occupants' safety is of great concern, 

two graphite/epoxy pressure hulls will be used rather than one. A material will separate the 

shells and increase the stiffness and strength without significant mass increase. Options 

included honeycombs of balsa, aluminum, or NOMEX, as well as urethane or PVC foams. 

The material most appropriate for this duty is NOMEX honeycomb core (Kander). NOMEX 

has high thermal tolerance, specific shear strength, and toughness. It is light weight and a 

good thermal insulator. These characteristics as well as its fire resistance also make it a 

prime candidate for the interior cabin. 

3.6 Flexible Material. The articulated frame dual-cylinder requires a flexible 

material to join the segments. A strong, highly flexible tube with a circular cross-section, 

having "accordion-like" characteristics is needed. The purpose of the flexible connection is 

to join the two structures, making essentially one continuous cabin. The rear structure is 

pulled by the articulated frame and not the flexible connection. An elastomer in the form 

of a thick solid, reinforced with KEVLAR 49 fibers, is a good candidate (Kander). One 

elastomer that may be used due to its favorable thermal properties is isobutylene isoprene 

(CRC Practical Handbook of Materials Science). However, lunar atmospheric effects on the 

material will have to be investigated. 

3.7 Graphite/Epoxy Selection. Merely stating that graphite/epoxy will be the 

material used for the structure is inadequate. Properties can vary quite drastically from one 

graphite/epoxy to another. Thornel, a subsidiary of Amoco, fabricates several candidates 

which have been used in space structural components. The tensile, compressive, and shear 

strengths as well as the modulus of elasticity should be high. The material should be 

thermally favorable, implying a very low coefficient of thermal expansion as well as a low 

thermal conductivity for insulation purposes. Finally, since mass is critical, the density 

should be minimized.

36
S

Design of a Pressurized Lunar Rover 24 April 1992 

A bar graph (Figure 3.3) was prepared comparing the different materials. Each 

material property was normalized against the best material in that category. To calculate the 

overall rating, an average of the ratings was taken. Each category was weighted equally and 

the best material was T-650/35/Epoxy. Its material properties are presented in Table 3.1 

S

(Thornel).

Table 3.1 - Properties of T-650/35/Epoxy 

Density 1,600 kg/m3 

Tensile Strength 2,068 MPa 

S

Modulus 152 GPa 

Thermal Conductivity 8.73 'W/(m.K) 

Expansion Coefficient -0.1 x 10-6/'F 

Compressive Strength 1,724 MPa 

Shear Strength 124 MPa 

S

S 

S

S

S

37 

0
Design of a Pressurized Lunar Rover 24 April 1992 

1.0 

0.8 

0.4 

0.2 

rol

• LVWUILi. I% 1⁄4JVI1ILL1!L '.%JUI1). alicar jveraii 

strength of thermal strength strength 

expansion 

P-55/Epoxy P- 100/Epoxy T-650/35/Epoxy 

P-75/Epoxy P-120/Epoxy T-650/45/Epoxy

Figure 3.3 - Graphite/Epoxy Candidates

38 
Design of a Pressurized Lunar Rover 24 April 1992 

3.8 Material Selection For Vehicle Substructure. The primary variable used in 

selecting a material for the saddle was weight. The lightest engineering metal, with the 

exceptions of magnesium and beryllium, is aluminum. Aluminum alloys are about one-third 

the weight of steel. They possess a high strength-to-weight ratio, are weldable, and can be 

easily machined. Thus an aluminum alloy will be used for the saddle. 

3.9 Conclusions. The materials selection and thicknesses for the rover's shell are 

presented in Table 3.2. Table 3.3 is the mass statement. Figure 3.4 is an illustration of the 

shell cross section. Structural analyses should be performed to check the validity of the 

materials and thicknesses chosen. 

Table 3.2 - Pressurized Lunar Rover Shell

MATERIAL DENSITY THICKNESS 

(kg/m3)

(cm) 

Graphite/Epoxy 1,600 0.25 

NOMEX Honeycomb Core 40 1.50 

Graphite/Epoxy 1,600 0.25 

Elastomer (Isobutylene Isoprene) 920 0.10 

MLI 1.21 (kg/m2) 3.00 

Aluminum 2,700 0.10

39 
Design of a Pressurized Lunar Rover 24 April 1992 

Table 3.3 - Materials Mass Statement 

COMPONENT MASS 

(kg) 

Graphite/Epoxy 553 

NOMEX Honeycomb Core 84 

Graphite/Epoxy 550 

Elastomer (Isobutylene Isoprene) 130 

ML! 130 

Aluminum 378 

Flexible connection 238 

Saddle 652

Elastomer 

Inner graphite/epoxy shell 

NOMEX honeycomb core

Aluminum 

Outer graphite/epoxy shell.

MLI 

Figure not drawn to scale 

Figure 3.4 - Pressure Hull Cross Section

40 
Design of a Pressurized Lunar Rover 24 April 1992 

4. MOBILITY SUBSYSTEM 

4.1 Configuration Selection. Numerous options were studied and compared before 

establishing a final mobility system configuration. The configuration options were broken 

down into frame, terrain interaction, and steering. 

4.1.1 Frame Options. As stated earlier, the frame option was a choice between a 

uniframe configuration and a multi-segmented one (Figure 2.3). Because of the extreme 

importance placed on vehicle mobility, a segmented vehicle was selected at the expense of 

added complexity. 

4.1.2 Terrain Interaction Options. Besides the conventional wheeled vehicle, tracks 

and walkers were evaluated. More exotic methods such as hoppers and rolling balls were 

even considered. Interestingly, prior to actual landings on the moon, the very same process 

took place among designers. It was believed that trafficability on the Moon's surface would 

be very poor. Consequently, many strange-looking walking machines were developed. 

However, from the experiences of the Apollo program, it is now known that any vehicle with 

round wheels will perform satisfactorily, provided the ground contact pressure is no greater 

than 7-10 kPa (Heiken et al). To be thorough, the more exotic schemes were still considered. 

They were not viable for our purposes. The walking vehicle concept was quickly rejected 

because of its complexity and its inability to meet the speed requirements. Much greater 

consideration was given to tracks or some variation thereof. The major advantage of a 

tracked vehicle is its increased maneuverability. Tracks provide better turn radii, traction, 

and slope climbing capabilities than do wheels. However, this increase in performance did 

not outweigh the drawbacks associated with its weight, complexity, and prohibitive power 

consumption for rover applications. Maintainability and reliability of the tracks in this 

environment was also a great concern. Thus a wheeled vehicle was deemed most appropriate. 

41
Design of a Pressurized Lunar Rover 24 April 1992 

Furthermore, it was decided that each wheel should be powered by its own electric motor 

placed within the wheel well. 

4.1.3 Steering Options. Three methods of steering were evaluated. The standard 

Ackermann system used by automobiles provided one possibility. The second option was to 

use an articulated frame steering system. The third possibility was to use skid (or scuff) 

steering commonly seen on tracked vehicles (Figure 4.1). 

The standard Ackermann configuration turns the front set of wheels, possibly via a 

rack and pinion arrangement. Another option is a double Ackermann system in which both 

front and rear wheels are turned. The standard Ackermann system was used as a baseline for 

comparison with the other steering systems. 

Since an articulated frame design was being considered, it made sense to have an 

articulated frame steering system option. This involves powering a yaw axis at the frame 

connection and rotating the vehicle sections with respect to each other. One of its primary 

advantages is that the wheels of the front and the back segments traverse the same paths. 

The driver knows that if the front segment clears an obstacle, the back one will. Also, the 

body can be bent to maneuver out of conditions that would stop a uniframe vehicle. 

Articulated frame steering produces turn radii smaller than the standard Ackermann system. 

A third option naturally evolved from the idea that each wheel should be individually 

powered by its own motor. Since each motor could be operated independently of the others 

and at different speeds, a skid steering system was considered. The skid steering concept 

involves driving the outer and inner sets of wheels at differential speeds (or directions). This 

scheme can generate turn radii much smaller than any of the other systems. However, this 

requires much more power than the other steering systems under consideration.

42
Design of a Pressurized Lunar Rover 24 April 1992 

Ackermann Articulated Steering Skid Steering Frame Steering 

Figure 4.1 Steering System Options

43
Design of a Pressurized Lunar Rover 24 April 1992 

4.2 Final Configuration. The overall mobility system and the various components 

of the system are described in the following sections. 

4.2.1 Saddle. Each pressure structure is supported by an aluminum saddle (Figure 

4.2), which is a cradle used to distribute the loads caused by the wheels. The saddle is made 

up of longitudinal and transverse support beams attached to the pressure hull as well as 

plating. The saddle angle for optimum load distribution is 168 0 (Moss). 

The two saddles are connected by an articulation joint (Figure 4.3). The joint 

controls yaw, and allows for free pitch and roll between the two pressure structures. The 

articulated turning of the rover is controlled by a motor attached to the yaw axis. The joint 

will be constructed of a high-strength steel. The joint will be subjected to stresses associated 

with maneuvering, but it will normally not have to pull the rear section since it will be 

driving with the same speed as the front section. Springs are place at the top of the flexible 

connection to balance the rover and keep it level. The joint will be housed in a flexible tube 

made of the same material as the flexible passageway. This will protect the joint from the 

effects of dust and thermal cycling. 

4.2.2 "Pinned Wheel" Concept. An unusual wheel concept was chosen for improved 

mobility (Figure 4.4). Each side of each segment of the vehicle is supported by two wheels. 

The wheels are connected at their centers by a hollow cylindrical bar which is pinned at its 

midpoint. A hollow cylindrical cross-section was chosen because it is lightweight, and allows 

electrical connections to be housed inside. The bar is attached at the pin through a passive 

suspension system to the saddle. Note that the bar's ends are bent away from the pressure 

structure making Ackermann steering possible. The bar is free to rotate about the pin. This 

allows a large degree of vertical movement by the wheels while still maintaining constant 

ground contact. Figure 4.5 shows the extreme conditions that the wheels can handle. Figure 

4.6 shows how the rover conforms to the terrain.

44
Articulatior 

powered y 

pitch and r 

Saddle distri 

loads on pre

Design of a Pressurized Lunar Rover 24 April 1992 

rngitudinal

All dimensions in meters. 

[

and transverse Saddle plating 4 meters support beams 

7.5 

Cylindrical bumper 

Figure 4.2 - Vehicle Substructure, Top View

45
Design of a Pressurized Lunar Rover 24 April 1992 

Ib

S

All dimensions in centimeters 

Pin diameter 2 c 

Side 

: 

15 15 I' 

25 

Figure 4.3 - Joint Schematic

46
Design of a Pressurized Lunar Rover 24 April 1992 

TOP VIEW Maximum tire SIDE VIEW 

rotation = 560 

Springs 

allow wheels 

to conform 

I

to terrain. 

Angular

05 motor provides 

torque to steer 

front wheel.

-

_;f0 

0.5 

'.-0.5--øj f 

0.5

0.2 

0.1 4 

0.65 

2.0 

Double 

transverse 

wishbone 

suspension 

connects wheel 

group to saddle

0.5 

LI

N6I 

All dimensions in meters.

I meter

Figure 4.4 - Wheel Group

47 

S 

S

0

S
Design of a Pressurized Lunar Rover 24 April 1992 

pj,1 

Fi O iI 

Figure 4.5 - Obstacle Extremes

48
Design of a Pressurized Lunar Rover 24 April 1992 

Figure 4.6 - Terrain Performance

49
Design of a Pressurized Lunar Rover 24 April 1992 

4.2.3 Suspension System. The suspension system connects the saddle to the "pin 

wheel" bars and connects the bars to the wheels. It is designed to position the wheels on the 

ground with a maximum contact area. Furthermore it transmits drive, braking, and turning 

forces, and spring and shock absorber loads. 

The proposed independent suspension system can be seen in Figures 4.4 and 4.7. It 

consists of a double transverse wishbone suspension (Figures 4.8 and 4.9) supporting the 

central pivot location of the "pinned wheel" arm. The upper and lower arms are of equal 

length to ensure that the connecting bar will not rotate about its longitudinal axis. A primary 

pair of components are two telescopic shock absorber-dampers balanced by springs. The 

purpose of the dampers is to dissipate some of the energy imparted to it in the form of heat. 

This leaves insufficient energy in the spring for violent rebounds, reducing the amplitude and 

frequency of oscillation. The upper end of the rod is pushed onto a ball thrust bearing 

housed in a rubber mounting. The bearing allows strut rotation while the rubber mounting 

accommodates angular movement and supports the body weight. The bottom end of the strut 

is pivoted to the lower arm by a ball and socket joint. Each suspension group is completely 

independent of the others, so the suspended body will not be influenced by small wheel

50
I0

Design of a Pressurized Lunar Rover 24 April 1992 

Ii 

- 

.-

N 

.-

51
Design of a Pressurized Lunar Rover 24 April 1992 

CJL31 PHO 'c H 

Figure 4.8 - Suspension Isometric

52
0 

Design of a Pressurized Lunar Rover 24 April 1992 

10.03 

Mounting 

stem and 

rubber 

bushes 0.45 

Telescoping 0.27 

shock 

absorber

10.1 N 0.08 

0.35 O.T.

Figure 4.9 - Double Wishbone Suspension

53 
Design of a Pressurized Lunar Rover 24 April 1992 

deflections (Heisler). Figure 4.10 shows the maximum suspension displacements that can be 

achieved while still maintaining a level cabin. A larger step would cause cabin rotation, as 

would prolonged exposure to these terrain conditions (ie. a long ditch). In the case of Figure 

4.10, the strut would force the wishbone members to return to their original rectangular 

shape, causing a counterclockwise rotation of the cabin. 

At the end of each bar, the wheel is attached via a spring which allows rotation about 

the longitudinal axis of the bar. Thus the wheels are allowed to conform to various surface 

conditions and ground contact is maximized. The springs are balanced so that the wheels 

remain vertical on a horizontal surface. 

4.2.4 Steering System. A dual steering system was chosen. Based on complexity and 

power consumption, it was decided that a combination of double Ackermann and articulated 

frame steering would be used. As the steering wheel is turned, an electronic control system 

turns the front and rear wheel pairs (Figure 4.11) (using motors mounted at the bar end) and 

the articulation joint to produce the appropriate geometry for an ideal, neutral steer turn 

(Figure 4.12). This geometry produces a minimum turn radius of 9.4 meters. However, after 

the articulated frame has been turned through its full angle, the ideal geometry may be 

sacrificed by turning the wheels further. This leads to a slight oversteer condition which will 

allow a tighter turn. In an oversteer, the vehicle will continue to turn while the steering 

wheel is held fixed. This will require that the driver be accustomed to controlling the vehicle 

during oversteer. A second option is to install an additional control system that can 

compensate for the oversteer. This insures that the response to steering control will be the 

same whether or not oversteer is present. 

In an emergency the driver has the option of utilizing skid steering, since the wheels 

are already individually powered. Although in itself this would be too power-consuming, it 

can supplement the primary steering system if required.

54
Design of a Pressurized Lunar Rover 24 April 1992 

55
0.5 meters 

Design of a Pressurized Lunar Rover 24 April 1992 

Figure 4.11 - Ackermann Steering System

56
Design of a Pressurized Lunar Rover 24 April 1992 

* Note: the actual vehicle turn radius can 

be further reduced by continuing to turn 

the front and rear wheels beyond the point 

of maximum frame rotation. However, 

the inner wheels will be unable to align 

producing an oversteer condition. 

Figure 4.12 - Ideal (Neutral Steer) Turn

57
Design of a Pressurized Lunar Rover 24 April 1992 

4.2.5 Wheel/Tire Assembly. The wheel/tire assembly Consists of the wheel well 

Containing the drive motor as well as the wheel, tire, and required gearing and support 

(Figure 4.13). The drive motor rotates a 0.45 meter radius rigid Cylindrical frame that is 

inside the tire and rigidly connected to the back of the wheel. Slightly elastic rings fill the 

area between the frame and the tire. These rings support the tire to enhance traction. The 

face of the wheel is covered to protect the drive system from large amounts of lunar regolith 

disturbing its operation. Grousers (cleats) will be used to enhance mobility system 

performance. 

4.2.5.1 Electric Motor. The drive motors and articulation motor of the PLR will have 

equal pull capabilities, so a single motor type may be used. Three-phase high performance 

DC brushless motors with variable speed drive will be used as the housed drive motors. 

These motors are mounted inside the hub of the wheel. The rated torque that can be 

produced is 900 Nm. Power at the rated torque is 1.5 kW. The maximum continuous stall 

torque provided will be a nominal value of 600 Nm (Inland). Smaller motors will be used for 

wheel turning, since they do not require such large torques. Hydraulic motors were not 

chosen due to the disadvantages they have for in-space applications. 

Brushless permanent magnet motors offer many advantages over conventional electric 

motors. Little maintenance is required with the brushless motor because it does not have the 

problems of brush longevity and maintenance. In addition, they have the capability of 

reversing torque, which greatly enhances maneuverability (Mateuffel). A three-phase power 

source system is preferable since it lowers wear on the motor (Turner). DC electric motors 

have a higher torque to weight ratio, are more reliable, and are easier to control (Jones). 

Thus a DC brushless electric motor was selected. 

Many advantages are found by using brushless systems with the proper servo 

amplifier. These include operation in harsh environments, a broad speed range, and many 

mechanical advantages. Figure 4.14 is a basic schematic of the system to be used. Hall

58
Sid 

Vi 

Flexible 

suspensic 

rings

Design of a Pressurized Lunar Rover 24 April 1992 

Top View 

Housed 

brushless 

motor 

Angular 

torque 

motor EM / 

71 Hybrid, single 

stage epicychc 

transmission 

meter 

Housed 

motor and 

transmission 

Figure 4.1-3 - Wheel Assembly

59
Design of a Pressurized Lunar Rover 24 April 1992 

1 

L

>

= 

= 

1= 

C-) 

E 

=

E 

cu 

- 

C > 

LZ

-

C

60 
Design of a Pressurized Lunar Rover 24 April 1992 

devices and current sensors are used for velocity and torque feedback to the amplifier. This 

feedback is compared with the input values and the servo amplifier adjusts the electrical 

output accordingly. This provides an overall reliable and rugged motor system. 

A compact transmission which utilizes multiroller planetary friction drives is used 

(Figure 4.15). A 50:1 speed reduction is possible to provide ample torque for difficult 

maneuvers. The transmission has a 92-93% efficiency (Nasvytis). 

Figure 4.15 - Multiroller Planetary Drive System

61
Design of a Pressurized Lunar Rover 24 April 1992 

4.3 Performance. The rover's mobility performance characteristics are presented in 

Table 2.2, reproduced here for convenience. 

Table 2.2 - Mobility Performance Characteristics 

Maximum Velocity 

Nominal Power Requirement 

Maximum Allowable Power 

Maximum Gross Pull 

Maximum Vehicle Acceleration 

Soil Compaction Resistance 

Normal Soil 

Soft Soil 

Hard Soil 

Minimum Turn Radius (in percent of vehicle length) 

Neutral Steer 

Oversteer 

Range 

Maximum Gradient 

Wheel Sinkage 

Normal Soil 

Soft Soil 

Hard Soil 

Climbable Step Height 

Crossable Crevice

(4 wheels), 14.7 km/hr

(8 wheels), 29.4 km/hr

(4 wheels) 1.36 kW/wheel

(8 wheels) 0.68 kW/wheel

1.5 kW/wheel

900 N/wheel

0.64 m/s2 

22 N/wheel 

81 N/wheel 

4.0 N/wheel 

78.3%

60%

2,000 km

26.5° 

1.03 cm 

2.86 cm 

0.21 cm 

0.53 M 

1.7 M 

A nominal velocity of 10 km/hr (2.78 m/s) was assumed, which will easily meet the 

mission requirements. Normally, the rover will be driven with only four of the eight wheels 

powered. The rationale is that if the rover becomes stuck while using all-wheel drive, it will 

have difficulty freeing itself. However, if the rover becomes stuck while using only half of 

the wheels, powering the remaining wheels will probably free it. 

Experimental analysis has shown that the coefficient of rolling resistance, f r

, is 

approximately 0.20 (Sponsler). The straight line power required is given by (Wong, ' J. Y.) 

Pet =

62
Design of a Pressurized Lunar Rover 24 April 1992 

• This must be scaled depending on the number of powered wheels. This means that the power 

required per wheel (using four wheels) for straight-line motion at the nominal speed is 1.36 

M With eight wheels powered, the power per wheel is 0.68 kW. The maximum power 

allocated to the drive motors is 1.5 kW. At this power, the rover can attain peak velocities 

.

of 14.7 km/hr and 29.4 km/hr, for four and eight powered wheels respectively. Turning 

power is difficult to estimate without experimental data for the specific steering system, 

terrain, and tires involved. It is estimated that the required turning power will be between 

0.25 kW and 1 kW, depending on the speed of the rover and the terrain. Low speed turning 

will require more power than high speed turning, so turning power will not affect the 

maximum power required for mobility. This will be defined by the fastest forward speed, 

• and the rover will not be traveling near this during maneuvers. 

Bekker's models of obstacle performance indicate that the rover should be able to 

climb steps 0.53 meters high and cross crevices 1.7 meters in width. Keep in mind that the 

wheels are two meters in diameter, so these values are not unreasonable. However, the step 

height is based on a stationary wheel, and the actual wheels should be able to far outperform 

standard wheels with regard to climbable step height. 

The following empirical formulas developed by Bekker (Heiken et al) were used to 

•

compute several of the performance parameters. The wheel sinkage, z, was calculated using

z (cm) = (W/Ak)1". 

The gross pull per wheel in newtons, H, was determined using 

H = (Acb + WtanØ,).[l_K/sL(1_eK)]. 

Figure 4.16 shows the variation of maximum gross pull per wheel with wheel slip. The 

operating envelope was chosen to allow maximum use of the soil's pulling capacity. This 

I defined the necessary torque output of the motor. The soil compaction resistance per wheel 

in newtons, R, was found using

R = [bk/(n+1)]z'1. 

Finally, the steepest slope a wheeled vehicle is capable of climbing may be estimated using 

63
Design of a Pressurized Lunar Rover 24 April 1992 

Slope = tan1[(H-R)/H]. 

The input parameters for these equations are 

W = wheel load (N) = 1,491 

A = wheel footprint area (cm') = 1,755 

L = wheel chord length of ground contact (cm) = 35.1 

b = wheel width of ground contact (cm) = 50 

k = cohesive modulus of soil deformation (N/cm 2) = 0.14 

k = frictional modulus of soil deformation (N/cm3) = 0.82 (normal soil) 

k = soil consistency (N/cm3) = ku/b + ko

= 1.8 

n = exponent of soil deformation = 1 (normal soil) 

cb = coefficient of soil wheel cohesion (N/cm 2) = 0.017 

Ob = soil/wheel friction angle (°) = 35 

K = coefficient of soil slip (cm) = 1.8 

s = wheel slip (dimensionless) = varies 

It has been suggested that some of these parameters be adjusted for the critical 

condition of soft soil (Wong et al). These isolated areas could immobilize the vehicle and thus 

must be taken into account. Therefore the following adjustments will be made. 

Soft soil k0 = 0.5 N/cm3 

Soft soil n = 0.5 

Hard soil k0 = 6 N/cm3 

Hard soil n = 1.25

64
S

Design of a Pressurized Lunar Rover 24 April 1992 

S

S

Pull per wheel in newtons 

S 1,000 I Maximum over-toroue condition 

ON 

S

600 Maximum rated motor capability 

OPERATING 

ENVELOPE 

0if

I I 

0 0.1 0.2 0.3

Wheel slip (dimensionless) 

S

Figure 4.16 - Pull Per Wheel Versus Wheel Slip 

S

01.1 

S

65
Design of a Pressurized Lunar Rover 24 April 1992 

5. POWER SUBSYSTEM 

5.1 Power Required. The first step in selecting a power system for the pressurized 

lunar rover was to determine the amount of power required. Power estimates were made for 

all systems on the rover and are listed in Table 2.4. With these estimates, several mission 

profiles were planned. The first profile was the Survey Mission. This mission was planned 

around a 14 day period, where 50% of the time would be inactive time (eating, sleeping, etc.), 

33% of the time would be survey work, and 17% of the time would be driving. The next 

profile was the Transport Mission. The profile was based on a four day period in which the 

vehicle was constantly driven. The third mission profile was the Search and Rescue Mission. 

This mission was based on a two day period where 80% of the mission will involve driving 

and searching and 20% of the mission involved the rescue. The power profiles for these three 

missions are shown in Figures 5.1 - 5.3. 

Based on the maximum average power of 8.125 kW from the transport mission profile, 

it was determined that a continuous power of 8 kW is needed. The driving factor in 

determining the peak power required was the mobility system because it draws the greatest 

amount of power. Normally the rover is driven by four wheels. In the event that all eight 

wheels are needed to be driven at once, the total required power output would be 12 kW. 

This established the peak power requirement. Based on these figures, the power system will 

consist of a primary power source creating 8 kW and a 4 kW storage capacity. 

5.2 Types of Power Systems. Several type of power generation were evaluated for 

the primary power source, including fuel cells, batteries, photovoltaic arrays, radioisotope 

thermoelectric generators (RTG), and dynamic isotope power systems (DIPS). Batteries were 

the first to be eliminated based on their short supply of power and long period of

66
Design of a Pressurized Lunar Rover 24 April 1992 

15-

- I I I I I I I I I I 

-

I I I I I I I

I I I I 

- I

I I I I I I

I I I I I I I 

- -------- -------------------- TQt&LA\ .E1Mi POWER: 3.92 KW' 

DRIVING PERIOD SURVEY PERIOD i SURVEY PERIOD , INACTIVE PERIOD 

0 I I I I I I I I I I I I I I I I I I I I I I if 

6 12 18 2 

TIME (HOURS) 

POWER REQUIREMENTS (W) 

PFRIC)fl 

ECLSS 

EXTERNAL 

FIXTURES 

INTERNAL 

REQUIREMENTS 

MOBILITY 

INACTIVE SURVEY DRIVING 

200 200 200 

0 1000 200 

1500 400 200 

0 0 6000 

300 600 600 

800 800 800

OPERATIONS 

S THERMAL 

CONTROL 

TOTAL: 2800 2m am 

S Figure 5.1 - Power Profile: Survey Mission 

(Typical 24 hour period)

67 

0
15 

S

7 

10 

Design of a Pressurized Lunar Rover 24 April 1992 

S

TIME (HOURS) 

POWER REQUIREMENT(W) 

ECLSS 200 

EXTERNAL 

ixmis 200 

INTERNAL 

REQUIREMENTS 200 

MOBILITY 6000 

OPERATIONS 600 

THERMAL 800 CONTROL 

TOTAL: 

Figure 5.2 Power Profile: Transport Mission

(Typical 24 hour period)

S

S

68
S

S

15 

10 

S

S

Design of a Pressurized Lunar Rover 24 April 1992 

S

TIME (HOURS) 

S

POWER REQUIREMENTS (W) 

SEARCH/ 

DRIVING RESCUE 

200 200 

400 1000 

200 800 

6000 0 

600 600 

800 800

8200 3400 

ECLSS 

S

EXTERNAL 

FIXTURES 

INTERNAL 

REQUIREMENTS 

MOBILITY 

OPERATIONS 

THERMAL 

S CONTROL

TOTAL:

Figure 5.3 - Power Profile: Search & Rescue Mission 

S (Typical 48 hour period)

69 

0
Design of a Pressurized Lunar Rover 24 April 1992 

recharging. The photovoltaic array was eliminated due to a large array area required. To 

generate 8 kW of power, a non-steerable array would require a surface area of in excess of 

80 m2. The next option eliminated was the fuel cells. Even though fuel cells could easily 

produce the 8 kW, the mass requirement for such a system would be approximately 3,000 kg 

(NASA PSS ESBD 91.1). RTG was disregarded because its low energy conversion efficiency 

(5-10%) and high mass requirement. For the required power, the mass estimate of the RTG 

is well over 3,000 kg (Elias). The last system was the DIPS. Most DIPS cycles were found 

to have an energy conversion efficiency between 20 and 30%, while their mass estimates 

where fairly low (500-1,000 kg). In terms of availability and technology, most DIPS were 

very mature. With these factors, the DIPS became the choice for the primary power system 

of the rover. For auxiliary power, batteries with a photovoltaic array for recharging were 

chosen because of their simplicity. 

5.3 DIPS. There are several DIPS cycles used in producing power. The first is the 

Brayton cycle. It uses a xenon-hydrogen mixture as its working medium. By using a gas, 

the need for a fluid management system is eliminated and fluid degradation does not become 

a problem (Stadnik). A second system is the Rankine cycle which uses toluene (a fluid) as 

its working medium. A third system is the Stirling cycle. The Stirling uses hydrogen as its 

working medium, again eliminating the need for a fluid management system. The energy 

conversion efficiency of the three system is: Brayton cycle 24%, Rankine cycle 21%, and 

Stirling cycle 40% (Walters). The technology for the Brayton and Rankine cycles is available 

now, while the Stirling cycle is still in the research and development phase (Walters). The 

Brayton cycle was chosen based on its working medium and availability. 

5.4 Closed Brayton Cycle (CBC). Figure 5.4 is a functional diagram of the CBC. 

The heat source units (HSU) of the CBC (Figure 5.5) are one meter long with a diameter of 

40 cm. Figure 5.6 shows the placement of the system within the pressure hull. They are 

70
Design of a Pressurized Lunar Rover 24 April 1992 

powered by the decay of Plutonium 238 modules, which have a diameter of 12 cm, length of 

5 cm, mass of 6.55 kg, and specific energy release (c1) of 0.9760 W/cm3. Each HSU contains 

17 modules. The thermal power output of each HSU, 4148 Wt, is calculated as follows: 

EH = 21w q• A L 

where At

 is the transverse area and L is the length of the rod. The CBC has a conversion 

efficiency of 24.1%. Therefore, in order to obtain 8 Me for the power system, an input of 

32.2 kWt is needed, based on the following calculation: 

E1 = E/efficiency 

In order to obtain this amount of thermal power, eight HSU's are needed. Since the mass of 

each HSU is 25 kg, the total mass of the eight HSU's will be 200 kg. 

The electrical power is created by the turboalternator compressor (TAC) assembly. 

The thermal power is transferred from the HSU to the TAC by the xenon-hydrogen working 

fluid. The gas is heated in the HSU and then enters the turbine end of the TAC. The 

temperature at the turbine is around 1150 K. The gas expands in the turbine and performs 

work, creating the electrical energy. The gas then passes through the recuperator where part 

of the excess energy is transferred to the cycle gas on its way to the HSU. The gas then leaves 

the recuperator and passes to the radiator were the excess heat is transferred directly to the 

radiator heat pipes, which then cycle the heat directly to space. The cooled gas is now at 

approximately 300 K. This gas travels from the radiator to the alternator (cooling it),

71
Design of a Pressurized Lunar Rover 24 April 1992 

I User Load I 

8 kW 

Heat

Turboalternator Compressor (TA C) 

Source Compressor i.iøl Alternator

Unit

----------------

300 K 

I Radiator I 

Recuperator 

Figure 5.4 - Closed Brayton Cycle Functional Diagram

72
MR Vacuum Liner>.< 

Multifoil Insulation (MFI) - 

Pin 

O

HX Outlet 

S

S

S

/ 

GPHS Module Canister

Him

RHRS 

Heat 

Pipe 

HX Inlet 

Design of a Pressurized Lunar Rover 24 April 1992 

S

S

S

Vacuum Li 

Variable Pitch HX 

S

Figure 5.5 - DIPS Power Module Heat Source 

S 

S

73 

0
0.0635 m 

Design of a Pressurized Lunar Rover 24 April 1992 

HSU 

40 cm diameter

Figure 5.6 - Placement of Primary Power System 74
S

Design of a Pressurized Lunar Rover 24 April 1992 

and then on to the compressor where the gas is compressed by a two to one ratio. From the 

compressor the gas travels again to the recuperator where it picks up excess heat from gas 

leaving the turbine. From the recuperator the gas travels back to the HSU and the cycle is 

complete (Johnson). 

The radiator is a very important part of the CBC. The total radiator area is 15 m2 

with a mass of 75 kg. The radiator itself is carbon-carbon fin with copper/water heat pipes 

running through it. The HSU's also have a thermal control system of their own for when the 

system is not at full operating capacity or a problem occurs in the working fluid cycle. This 

system only operates when the HSU assembly temperature rises above 1300 K. The heat 

pipes are molybdenum with a rhenium coating and lithium as the working fluid (Schmitz). 

5 Shielding is required in order to protect the crew from the radiation given off by the HSU.

The shield will be made of tantalum alloy, which will partially encase the HSU. 

The volume breakdown of the CBC is as follows: 

HSU 1.006 m3 

TAC 0.5 m3 

S

Shielding 0.4818 m3 

Total volume 1.9878 m3 

0 5.5 Secondary Energy Storage. A secondary energy storage system is needed to 

handle the peaks requirements above 8 Me. The maximum peak power is 12 kW, which is 

4 kW above the power put out by the CBC. It was estimated that these peaks would last no 

longer than an hour at a time. From the power profiles, the needed auxiliary power was 

estimated to be 4 kWhr. To satisfy this need, a sodium-sulfur battery system was chosen for 

the secondary power system, with a photovoltaic array to recharge the batteries. The Na/S 

75
Design of a Pressurized Lunar Rover 24 April 1992 

battery has a specific power of 250 Whr/kg (Withrow). With these figures, the total mass of 

the batteries would be: 

mass of batteries = storage energy / specific power = 16.0 kg. 

Once the batteries have been partially discharged, they can be recharged in one of two ways. 

The first is by the photovoltaic array made of GaAs/Ge. The second is by the primary power 

source, the CBC. When the total energy from the CBC is not being used, some of it can be 

redirected to charge the batteries. The size of the photovoltaic array was chosen to be 9 m2. 

With an efficiency of 18.3%, the total power collected by the array will be 2.0159 kW. With 

this power, the batteries will be charged at a voltage of 240 V with a current of 8.4 amps. 

When the CBC is used to charge the batteries the voltage will still be 240 V but the current 

will vary according to the available power at the time. 

5.6 Power Management and Distribution (PMAD). The purpose of the PMAD 

(Figure 5.7) is to distribute and convert the electrical energy produced by the CBC. The 

PMAD is made up of all the electrical components outside the CBC. The CBC produces 8 

kWe of single phase AC power at around 200 volts. This power is regulated by a shunt 

regulator which dissipates all power not being used by the load. The power conditioning is 

done at the loads through converters designed for each (Schmitz).

76
Design of a Pressurized Lunar Rover 24 April 1992 

LOAD DISTRIBUTION CENTER 

I----------------- I

I 

EMS ________ I t converter i .5 KW/whee

I 

I I

I I 

I

I 

I 1 

EXTERNAL

I 

I

I 

FIXTURES i converter i

I I 

I

I 

I I I

I I I 

INTERNAL

I I

I I 

REQUIREMENTS i converter

I

I I I $ I I I

I I I I

I I I I

I I 

MOBILITY

I 

I converter I

I 

I I

I I I

I 

OPERATIONS

I 

I ______ I I converter Li_

I I 

I

I I $

I I I I

I I I I

I I 

THERMAL

I

I 

CONTROL I 1 I

-

SHUNT

POWER 

SUPPLY 

Figure 5.7 - Power Management and Distribution Functional Diagram

77 
Design of a Pressurized Lunar Rover 24 April 1992 

6. THERMAL CONTROL SUBSYSTEM 

The temperature of the PLR is controlled by the balance of three thermal entities. 

1)heat absorbed from thermal radiation (solar and lunar) 

2) heat generated within the spacecraft 

3) heat lost by radiation through the exterior surfaces 

The first and third entities are considered the most important in design; however, the second 

cannot be ignored. Internal heat accumulation could pose a serious problem to a manned 

space vehicle. Common internal sources of heat are: 

• waste heat and heat losses from power source 

• electrical power and subsequent dissipation of heat by electrical resistance and 

mechanical friction of moving machinery 

• human metabolism 

• chemical processes 

A functional diagram of the thermal control system is provided in Figure 6.1. 

6.1 Options. There are basically three types of thermal control systems. The first is 

a passive system. In a passive system, there are usually no moving parts or heaters. The 

second type is a semi-passive system. It differs from the passive system because there are 

simple temperature-activated controls to close or open conductive paths, transport heat 

through heat pipes, and control the direction of radiation. Finally, there are active systems. 

They include pumped-loop systems, heaters controlled by thermostats, and mechanical 

refrigerators. Examples of all three are shown schematically in Figure 6.2.

78
Design of a Pressurized Lunar Rover 24 April 1992 

S

Radiator 

S _with Active 

Louvres 

Electronic I J

Heat I 

Equipment pipes

Radiator Human p 

etabolis 

Power 

S Source 

L

Solar 

Radiation 

Multilayer Heat 

Insulation Pipes Lunar 

S Radiation 

Climate 

Control white 

Epoxy 

S Coating 

Multilayer 

Insulation 

S

Figure 6.1 - Thermal Control Functional Diagram 

S

S

79 

S
S

Pump 

'I

Design of a Pressurized Lunar Rover 24 April 1992 

S Passive: __________ Space Radiator 

Cold Plate 

Phase Change

Device 

Electrical 

Equipment 

S

Semi-Passive: Thermal Control Louvres

Space Radiator 

Cold Plate 

Phase Change

Device 

Electrical 

Equipment 

S Active: 

S

Figure 6.2 - Thermal Control Options (Wertz et al.)

80 

S
S

Design of a Pressurized Lunar Rover 24 April 1992 

A passive design is generally lighter, requires less electrical power, and costs less than 

an active design. Active designs are needed when there are small tolerances for temperature 

control. A reasonable design approach is to begin with a passive system and add additional 

elements as needed (Van Vliet). 

I

6.2 Internal Thermal Control. Considering the size of the PLR, and the nature and 

time of a typical mission, all three of the basic thermal control systems will be used. Some 

0 temperature tolerances for the PLR are presented in Table 6.1 (Fortescue et al). 

O

Table 6.1 - Temperature Tolerances 

Electronic equipment (operating) -10 to +40°C 

Batteries -5 to +40°C 

Microprocessors -5 to +40°C 

Bearing mechanisms -45 to +65°C 

Solar cells -45 to +65°C 

Humans +18 to +30°C 

Internal heat dissipation will also need to be taken into consideration. Some sources of heat 

rejection and their amounts are (Walters, Eagle Engineering): 

Human metabolism (4 men) 0.342 kW 

Electronic equipment 3.0 kW 

Power source 25 kW 

6.2.1 Electrical Equipment. Most of the electrical equipment is located in two

centralized areas; near the driving stations in the front module and the work stations in the

rear module. In each of these two locations, a small, independent looping system will be set

up. In both cases, heat will be transported from the equipment to the space radiator. The

81
Electrical

Equipment

#3 

I Cold Plate I

Electrical 

Equipment 

I Cold Plate I 

S

Design of a Pressurized Lunar Rover 24 April 1992 

radiator itself will be a second surface mirror (SSM). The heat transported to the SSM will 

S be immediately rejected into space. A total of 1.5 m2 of radiator area will be needed for the 

electrical equipment. A radiator area of 1.0 m 2 will be used for the equipment at the drive 

station and a 0.5 m2 radiator will be required at the work stations. A generic functional 

S diagram of this loop is given in Figure 6.3. 

0

S

Heat 

Pipe 

S

Electrical Electrical 

Equipment # Equipment 

1 #2 

Cold Plate Cold Plate 

Figure 6.3 - Thermal Loop For Electrical Equipment 

S

S 

0

82 

0
n

S

Design of a Pressurized Lunar Rover 24 April 1992 

S 

Thermally activated louvres will be placed on top of the SSM's. When the louvres are 

open, the radiator will function normally, but when closed, a heat shield will be formed. The 

rationale is that when the temperature is colder outside, the heat dissipated from the electrical 

equipment will stay inside the rover. Therefore less stress will be place upon the climate 

S control system. The angle at which the louvres will be oriented will be linearly related to the 

external temperature. A schematic of this design as well as a control block diagram is shown 

in Figures 6.4 and 6.5 respectively. 

.

S

Figure 6.4 - Radiator with Active Louvres 

83 

.

0
Design of a Pressurized Lunar Rover 24 April 1992 

S

S

S

External Voltage Moment 0 

Temperature T1 h

01/ 0 g 01 

Figure 6.5 - Control Block Diagram of the Radiator

with Active Louvres 

6.2.2 Power System. The most critical part of the interior thermal design will be the 

rejection of the heat produced by the power system. The PLR will use a Closed Brayton 

Cycle. This system rejects approximately 25 kW of waste heat (Johnson). The power system 

will be located in the rear module (under the floor and near the back). The power system 

will need to be insulated. Thirty shields of multilayer insulation with a conductivity of 

0.00029 W/mK would serve this purpose (Wertz et al).

84
Design of a Pressurized Lunar Rover 24 April 1992 

Ideally, the heat should be rejected immediately to a space radiator from the power 

source, and the power source should be located as high as possible. However this conflicts 

with the vehicle's stability requirements. The radiator surface cannot be put on the 

underbelly of the PLR because the heat flow of the lunar surface would degrade its 

effectiveness. Also, the saddle for the mobility system would interfere, therefore that option 

was dropped. Alternatively, the radiator could be put on top of the rover. The only 

drawback of this scheme is that an elaborate heat pipe system is needed to transport the heat 

from the power system near the bottom of the rover to the space radiator on top. 

The size of the radiator may be calculated using the Stefan-Boltzmann relation, 

Q/A = ac(T1 4 - T24) 

where 

Q = heat rejection = 25 kW 

A = radiator area (unknown) 

a = Stefan-Boltzmann constant = 5.67x10 12 W/(cm2K) 

= emissivity of radiator = 0.85 

Ti

 = temperature of radiator (source) = 500 K 

T2 = external temperature (sink) = 227 K 

From this calculation, a radiator with a 15 m2 surface area will be needed to reject the heat 

from the power source. 

6.2.3 Heat Pipes. The thermal control system will be required to reject the 25 kW 

of waste heat produced by the Closed Brayton Cycle. A critical factor in the heat rejection 

is that the heat must be transported over a three meter distance. This can be accomplished 

with heat pipes.

85
Design of a Pressurized Lunar Rover 24 April 1992 

A heat pipe is a closed tube whose inner face is lined with a porous capillary wick. 

The wick is saturated with the liquid phase of the working fluid and the remaining volume 

of the tube contains the vapor phase. The heat applied at the evaporator (at the heat source) 

vaporizes the working fluid in that section. The resulting difference in pressure drives vapor 

from the evaporator to the condenser (at the space radiator) where it condenses, releasing 

latent heat of vaporization to a heat sink in that section of the pipe. Depletion of liquid by 

evaporation causes the liquid-vapor interface in the evaporator to enter into the wick surface 

and a capillary pressure is developed there. This pressure forces the liquid back to the 

evaporator for re-evaporation. This process carries on continuously. 

The operating temperature of the heat pipes is relatively high (600 K). For such a 

high temperature, H20 would be a good candidate for the working fluid. A compatible 

material for the container and wick must be selected for the H20 working fluid. If a poor 

choice is made, harmful effects, such as corrosion and chemical reactions, may occur. 

Copper is highly suitable for the H 20 working fluid (Chi). 

When designing a heat pipe, there are several limitations that must be taken into 

account. To prevent turbulent flow as well as compressibility effects, the flow must be less 

than Mach 0.2. This speed limitation yields a sonic limitation on the heat flux (Q5). 

There is also an entrainment limit on the heat flux (Qe) Since the vapor and the 

liquid move in opposite directions in a heat pipe, a shear force exists at the liquid-vapor 

interface. If the vapor velocity is sufficiently high, a limit can be reached at which liquid 

will be entrained in the vapor. Once entrainment begins, a substantial increase in fluid 

circulation occurs until the liquid return system cannot accommodate the increased flow. 

This eventually leads to an abrupt dryout of the wick at the evaporator. 

Another limiting factor on the heat flux is the capillary pressure (QL). For proper 

functioning, the pressure at the liquid side of the liquid-vapor interface must be different 

from that of the vapor side along the entire length of the heat pipe. The pressure difference 

is called the capillary pressure. If this pressure is lost, the circulation will cease.

86
Design of a Pressurized Lunar Rover 24 April 1992 

The final limiting factor for the maximum heat flux is the boiling temperature (Q1,). 

The boiling limitation is basically a limitation of the radial heat flux at the evaporator and 

condenser. The limitation is set so that no vapor forms in the wick. If vapor bubbles did 

form in the wick, hot spots would be created which would obstruct the circulation of the 

liquid.

The limiting heat fluxes should be greater than the expected heat flux performance 

of the pipe. Following the graphical/analytical procedure for the design of heat pipes in S. 

W. Chi's, Heat Pipe Theory and Practice, heat pipes were designed for the transportation of 

waste heat from the power source to the radiator. Specifically, 20 heat pipes will be used to 

transport 1.25 kW of heat each. The heat pipe configuration is shown in Figure 6.6. 

6.2.4 Internal Air Temperature. There are basically two situations to design for when 

considering the internal air temperature. From the functional diagram in Figure 6.1, lunar 

and solar radiation will be a factor during a lunar day. Large amounts of thermal radiation 

will penetrate the PLR. A climate control system (i.e a heat pump) will be required to 

control the temperature. 

In the front Driving/Habitation Module, an intake fan will be located above the 

driving station. This will draw in the air and a temperature control device will correct the 

temperature. Then the air will be blown back to the rear of the cabin. A similar scheme will 

be utilized in the rear Work/EVA Module. In that module, the air will be taken in above the 

exercise station and put back into the cabin in front. 

During a lunar night, there will be no solar radiation. Instead, the internal heat of the 

PLR will radiate outward. To design for this situation, a heat piping system emanating from 

the power system will be set up to transport waste heat to the rest of the rover. A thermal 

switch will be employed to open or close the conductive path of the pipes depending on the 

situation. The climate control system will still be utilized to guarantee the small temperature 

tolerances.

87
Heat In-Evaporation Wick Heat Out-Condensation 

Copper pipe with copper wick 

with water working fluid HEAT FLUX LIMITATIONS 

SeId Pipe or 

opper) 

Design of a Pressurized Lunar Rover 24 April 1992 

outside diameter 2.54 cm 

inside diameter 2.21 cm 

vapor core diameter 1.02 cm 

wick thickness 1. 19 cm 

end cap thickness 0.33 cm

copper wire screen wick 

Mesh #300

Sonic Q s -586 kW > 1.25 kW 

Capillary Pressure QLc 1.4 kW> 1.25 kW 

Entrainment Q e 58.3 kW> 1.25 kW 

Boiling Qb 1.32 kW> 1.25 kW 

Figure 6.6 - Heat Pipe Schematic

88 
Design of a Pressurized Lunar Rover 24 April 1992 

6.3 External PLR Thermal Control. Most thermal spacecraft designs advocate the 

practice of completely insulating the craft from the external heat environment while 

controlling the rejection of heat in the interior of the craft (Van Vliet). 

To calculate the heat flux through the PLR shell, the equilibrium skin temperature 

is used (Fortescue et al).

A PLR ECT4 = A5aJ5 + AM aJA + AMEF12JN

where 

APIR = effective radiating area of the PLR

c = emissivity of the PLR skin = 0.85 

a = Stephan-Boltzmann constant = 5.67 x 10-8 W/m2 

T = equilibrium skin temperature (K) 

As = projected area of PLR towards the sun (m2) 

a = absorptivity = 0.2 

is = solar radiation = 1400 W/m2 

AM = projected area of PLR towards the Moon (m2) 

= solar radiation due to lunar surface albedo (98 W/m2) 

F12 = view factor 

JM = lunar radiation = 1120 W/m2 

During the lunar day, the following estimates were used:

APLR = 81.7 m2 

As= 49.02 m2 

AM = 44 m2 

F12 = 0.6

89 
Design of a Pressurized Lunar Rover 24 April 1992 

The calculation resulted in a skin temperature of 346 K. During the lunar night, the effects 

of the direct solar radiation as well as the reflected solar radiation are not present. The 

effective radiating area was increased to 163.4 m2, the view factor was estimated to be 0.5, 

and AM remained the same. The skin temperature for the lunar night was calculated to be 

227K.

The thermal conductance through the shell is found using: 

S q = 27rL(T ,..r - Touted/ (E [ln(r 1 /r1.1 )1/k 1 ) (i = 2,3,...7) 

Inner G/E r2 = 4.0025 m k = 8.73 W/mK 

r

i

 = 4.0000 m 

NOMEX honeycomb r3 = 4.0175 m k = 0.03 W/mK 

core r2 = 4.0025 m 

Outer G/E r4 = 4.0200 m k = 8.73 W/mK 

= 4.0175 m 

S Elastomer r5 = 4.0210 m k = 0.035 W/mK 

= 4.0200 m 

ML! r6 = 4.0718 m k = 0.00029 W/mK 

= 4.0210 m 

S

Aluminum bumper r7 = 4.1128 m k = 200 W/mK 

shield r6=4.1118m 

Overall Shell k = 0.160 W/mK 

The heat fluxes for extreme external conditions were found to be: 

Maximum heat flux during lunar day = 8.30 W/m2 

S Maximum heat flux during lunar night = 10.69 W/m2

S

90 
Design of a Pressurized Lunar Rover 24 April 1992 

6.3.1 Multilayer Insulation. Multilayer insulation (ML!) consists of layers of 

aluminized Mylar alternated with coarse net material (Wertz et al). The net acts as a separator 

for the Mylar layers, thus trapping air. Usually, Kapton is used for the outermost and 

innermost layers. ML! has a thermal conductivity ranging from 0.00029-0.0004 W/mK 

depending upon the number of shields. A three centimeter thick layer of ML! will be used 

to insulate the structure from the external lunar and solar radiation. 

ML! serves two other purposes. First it increases the effective heat capacity of the 

vehicle, reducing the harmful effects of thermal cycling. Second its light weight will reduce 

the need for highly active thermal control systems, thus increasing reliability with reductions 

in cost and mass. 

6.3.2 White Epoxy Paint. The whole exterior surface of the shell will be coated with 

a white epoxy paint. The paint will not affect the thermal absorption of the surface, but 

will increase the emissivity. This results in a considerably lower a/c ratio which can greatly 

reduce the temperature of the skin. For the paint chosen, if the exterior temperature were 

422 K, due to the low a/c (0.28), the skin equilibrium temperature would be 287 K (Wertz 

et al).

6.3.3 Windows. Special coatings will be necessary for the windows. To prevent the 

penetration of UV light, 0.1% Iron Oxide (Fe202) must be fused into the glass (Van Vliet). 

When sunlight transmits through glass, there is an internal temperature increase associated 

with it. This situation is more serious when dealing with windows on the moon. A thin 

film of gold will need to be sprayed onto the exterior surface of the window (Van Vliet). 

Heat penetration is considerably reduced while the window remains transparent. 

6.3.4 Solar Cells. The effectiveness of solar cells diminishes with increasing 

temperature (Fortescue, et al). Also an appreciable amount of the solar radiation that hits 

91
Design of a Pressurized Lunar Rover 24 April 1992 

the cells are unusable because it is not within the range of wavelengths suitable for power 

conversion. This waste radiation only heats up the cells and degrades the effectiveness. A 

solution to the thermal problem is to coat the cells with silica. The silica increases the 

emittance, thus lowering the skin temperature. However, there is a 2% reduction in energy 

available for power conversion. To remedy the problem with the unusable radiation, filters 

will be used to block out the bad radiation, thus lowering the absorptance by as much as 50%. 

This will reduce the power output of the solar cells by slightly less than 10%, but the 

S effectiveness will increase. 

Table 6.2 - Thermal Control Subsystem Mass Statement 

S

Component Mass 

(kg) 

Radiators 90 5 ML! 130 

Heat pipes 100 

Heat exchanger 50 

Pump 20 

Refrigerant 50

S

S

S

S

92 
Design of a Pressurized Lunar Rover 24 April 1992 

7. LIFE SUPPORT SUBSYSTEM 

The purpose of the life support subsystem (LSS) is to maintain proper conditions for 

the crew and equipment inside the rover. 

The first step in establishing the life support subsystem was to determine the design 

loads. These are shown in Table 7.1. Then the characteristics of the subsystem required for 

these loads were determined and shown in Table 7.2. 

Table 7.1 - Life Support Subsystem Design Loads

Metabolic 02 0.83 kg/man day 

Metabolic CO2 1.00 kg/man day 

Potable H20 1.86 kg/man day 

Metabolic H2O 

Handwash HO

0.35 kg/man day 

1.81 kg/man day 

Shower H20 3.63 kg/man day 

Perspiration and Respiration H20 1.82 kg/man day 

Urinal Flush H20 0.49 kg/man day 

Urine H20 1.50 kg/man day 

Food Solids 0.73 kg/man day 

Food H20 0.45 kg/man day 

Food Packaging 0.45 kg/man day 

Fecal Solids 0.03 kg/man day 

Charcoal required 0.06 kg/man day 

Trash 0.82 kg/man day 

Trash Volume 0.0028 m3/man day

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Design of a Pressurized Lunar Rover 24 April 1992 

Table 7.2 -. Life Support Subsystem Characteristics 

Subsystem Mass (kg) Power (W) 

Temperature and Humidity Control 50 60 

Atmospheric Control and Supply 80 10 

Atmospheric Revitalization 20 10 

Fire Detection and Support 10 --

Water Recovery and Management 120 30 

Waste Management 20 20 

Extra-Vehicular Activity & Support 20 30 

Total 320 160

The next step was to establish the degree of loop closure, i.e. how many of the 

expendable resources, if any, should be reclaimed. After examining the available technology 

as well as the weight and power penalties involved, a partially closed system was chosen. A 

completely closed loop system for such a small scale does not exist. Furthermore, unlike an 

open system, a single unit failure would jeopardize the entire closed loop. Since the mission 

is only 14 days, only the water and air will be reclaimed. The technology of the future 

should be able to perform this duty without a major power or weight penalty. The LSS 

operates under the parameters listed in Table 7.3. 

Note that the total pressure of the cabin varies from 68.3 to 101.325 kPa. The cabin 

will be pressurized to 101.325 kPa when docking with the lunar base which will also be 

pressurized to 101.325 kPa. When departing from the lunar base, the cabin will be 

depressurized to 68.3 kPa so that the prebreathing before EVA is minimized.

94 
Design of a Pressurized Lunar Rover 24 April 1992 

Table 7.3 - Life Support Subsystem Parameters

Parameter Nominal Conditions 

CO Pressure 

hessure

33.3 Pa 

02 18.6-22.06 kPa 

Total Pressure 101.325 kPa 

Temperature 65-75 OF 

Dew Point 40-60 OF 

Ventilation 4.6-12.2 rn/mm 

Potable Water 3.08-3.67 kg/man-day 

Hygiene Water 5.44 kg/man-day Wash Water 12.7 kg/man-day

The LSS functional schematic is shown in Figure 7.1. The carbon dioxide is removed 

from the cabin air and then the carbon is separated from the oxygen through reduction. 

Meanwhile, the waste water undergoes oxygen generation where the hydrogen and oxygen 

are separated by electrolysis. The oxygen is sent back to storage while the hydrogen is 

combined with the oxygen from the CO2 reduction, and water condensate is stored and 

recovered. The water is then used for non-potable functions. Thus the air and water are 

both reclaimed.

95 
Design of a Pressurized Lunar Rover 24 April 1992 

Vent and 

Relief

Oxygen storage 

Airlock

Fecal collection, 

processing, and 

Nitrogen storage 

/Oxygen I 

THC

Air cont. L Nit Hygiene station' rogen 

S

storage 

FDS Q 02 N2

Air Q asi Waste water 

CABIN I

CO2 -KD--' I 02 Shower 

Products Water i1ii•I wash 

•1

rag

Heater 

Co2 CO2 Cond. Cond. moval reduct. stor. recov.

Potable water 

Oxygen 

generation Water 

Figure 7.1 - Life Support Subsystem Functional Diagram

96
Design of a Pressurized Lunar Rover 24 April 1992

8. GUIDANCE, NAVIGATION, AND CONTROL SUBSYSTEM 

8.1 Guidance and Navigation. The rover's guidance and navigation requirements will 

be met by an inertial guidance system and the use of satellite observations. The satellite will 

map areas in which the rover will operate. The inertial guidance system will overlay the 

position coordinates so that relative positions may be known and seen from a top-view. 

The inertial guidance system is shown in Figure 8.1. The ball bearings on all the axes 

allow the gimbals to rotate with respect to each other and the case itself. The inertial 

guidance system is set with respect to the local horizon coordinate system at position 

(x0,y0,z0,t0). The gyroscopes are used to keep the gimbals in that position during the mission. 

When a gyroscope detects a disturbance tilt, it generates an electrical response which is 

processed and then directed to a gimbal motor. This motor applies a torque to rotate the 

appropriate gimbal in the proper direction to remove the tilt. The gyroscopes are placed 

orthogonal to each other so a disturbance in any axis may be discovered and the system may 

be kept stable. 

The accelerometers are also placed orthogonal to each other. With the use of a spring 

scale, the accelerometers are able to detect the rover's acceleration. Integrating yields the 

velocity and position (relative to the starting point). The navigation computer will keep track 

of the rover's movements and overlay the coordinates on a satellite map of the region. 

The gyroscopes sometimes produce errors in position coordinates when acted on by 

torques. Therefore an optical tracker will be used to provide redundancy and safety. Two 

celestial objects, a substantial angular distance apart, are necessary to keep the three-axis 

gimbal system stable. Ideally, three trackers should be used, so adjustments to all three axes 

may be made simultaneously and the system can pick up a celestial body entering the view 

while another celestial body is leaving. However, only one tracker is used because of their 

bulkiness. It is switched alternatively to the selected celestial bodies. Figure 8.2 shows the 

optical tracker. The functions of sensing the celestial body, adjusting the telescope into

97
6ACCLLLkCMLTLRS SLNS.a. ALOt.L .MI!AL 

ER 

Figure 8.1 — Inertial Guidance System

(as taken from Slater) 

Figure 8.2 — Optical Tracker

(as taken from Slater)

98
Design of a Pressurized Lunar Rover 24 April 1992 

alignment with it, and calculating the angles are all done by the on-board computer, with the 

use of a photosensitive device placed behind the focal plane of the telescope. 

The driver will be able to see any obstacle within three meters of the rover through the 

front windows. Also lights and cameras are placed in strategic locations to see any blind 

spots. To help the driver determine if a rock is "clearable", a laser will be positioned on the 

front of the rover. When on relatively flat land, if the driver sees red light being reflected 

off of a rock, he will know that the rock is not clearable. Another problem in navigating will 

be to clear rocks when going over hills and large bumps. For example, while the rover is 

ascending a hill and the rear cylinder has not reached the hill yet, the pinned bar will tend 

to rotate the front of the rover down, decreasing the ground clearance at the front end of the 

rover. In case of a rock obstacle, a cylindrical bumper/roller made of elastomer will be 

placed in front of the rover. If a rock hits it, it will turn the front part of the rover about 

the pin. This will help the rover clear the rock and at the same time protect the underbody. 

8.2 Control. A manual system is used to control the rover. The steering wheel will 

control the front two wheels (of the forward cylinder) and the rear two wheels (of the rear 

cylinder). It will be supported with the articulated joint, with skid steering available in an 

emergency. The driving mechanism will consist of three pre-programmed modes. These 

modes will depend on how many powered wheels are needed (i.e. four, six, and eight 

powered wheels). Also there will be a manual override capability with eight individual 

controls to power each wheel. 

While steering, the computer will calculate all the necessary angles for a turn of 

perfect geometry, using the articulated joint and Ackermann steering until the maximum 

articulation angle is achieved. If a tighter turn is required, it can be achieved using oversteer 

(see §4).

99
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Design of a Pressurized Lunar Rover 24 April 1992 

0

9. COMMUNICATIONS SUBSYSTEM 

The missions of the rover require several different communications links. These links 

will be combined into one subcarrier scheme, amplified by a common power amplifier and 

S

broadcast by one antenna. 

Earth-to-lunar links are nearly constant in range (about 385,000 kilometers). Due to 

the effects of lunar radio noise, atmospheric absorptions, etc., the equivalent noise 

temperature of 400 K is assumed. 

It is assumed that there will be a network of three lunar stations which will allow full 

coverage of the moon for voice transmission (Filipowsky). 

9.1 The S-Band System. The communications system requires a ranging system for 

accurate range measurements. This is accomplished by the repeated transmission of a long 

pseudo-random code. The code bandwidths are very wide, so they must be broadcast in the 

C-band or the S-band. 

Also, there are two approaches to dealing with the separate links of the 

communications system. The VHF approach involves transmitting each individual link with 

S

its own antenna. These frequencies are allocated in the UHF and VHF regions. The S-band 

approach combines the separate signals into one signal, whose wide bandwidth requires 

allocation in the S-band. This eliminates the need for multiple antennas. 

By combining the ranging system with the unified approach, the communications 

subsystem becomes more compact and requires less power. The various signals are combined 

with a modulator, then amplified by a 20 W amplifier for transmission at 2.3 GHz by a 

• parabolic reflector antenna. The signal can then be transmitted to either a lunar base, or 

directly to earth. 

S

100 

0
Design of a Pressurized Lunar Rover 24 April 1992 

9.2 Voice Links. The voice link (Figure 9.1) allows direct communications between 

the astronauts and the lunar base, with the option of communication with earth, if necessary. 

For redundancy, two separate two-way voice links are provided. It has been shown that a 

three kHz band transmitted at 30 dB will produce easily understandable speech (telephone 

S

lines use the same bandwidth, but at a higher signal contrast). The range equation can be 

used to determine the transmission power. 

Pr = kTopBtrO147TR2M/GAT7 

where 

k = Boltzmann constant 

T, = average operating radio noise temperature 

Btr = transmission bandwidth 

a 1 = input signal contrast 

R = distance from source 

M = safety factor 

G = gain of broadcasting antenna 

A = area of receiving antenna 

= antenna efficiency 

S

By assuming a safety factor of 10, a 26 dB gain, and an 85 ft diameter receiving 

S 

antenna with an efficiency of 60%, the voice link will require 1.32 W of power. 

The voice link will allow normal conversation between the astronauts and the base, 

with a 2.6 second round trip delay between the rover and earth, requiring a slightly altered 

conversation scheme. Also, a special frequency is allotted for emergency messages. 

The rover is equipped with a voice control center which will also serve as an 

intercom. It accepts inputs from standard microphones located throughout the rover and 

provides outputs to loudspeakers. The power requirement for voice transmission is about 2W, 

which may need to be modified depending on ground equipment.

101 

S
S

Design of a Pressurized Lunar Rover 24 April 1992 

To and SC Antenna 

from ____________________

astronaut _______________________ V 

• ____ I 

I Voice control center I I Voice transceiver

sc Antenna 

5

L_.J Alternate transceiver 

S Figure 9.1 - Two-way Voice Link 

9.3 Telemetry Link. The telemetry link (Figure 9.2) transmits all data information 

from the rover to the base. This data includes scientific data (from measurements), 

biomedical data, housekeeping and structural data, and computer processed data. This data 

• will be transmitted in real time when the rover is in line-of-sight with the base. For 

long-range missions, the telemetry can be stored onboard and transmitted at high speed when 

near the base. The telemetry will be processed at the base, then transmitted to earth. 

• Telemetry data will be handled with PCM transmission, a type of digital transmission 

that is very flexible and allows a high degree of accuracy. The telemetry will also carry some 

analog data. The analog data will be combined with the digital data to produce a composite 

signal which is then transmitted. 

S

The signal power contrast will be about 13 dB (which will result in an error rate of 

about I in 10), and the signal will transmit about 108 bits/day, or 2500 bits/sec. 

S

102 

41
From 

sensors

Onboard 

processor Combiner I-,.I Transmitter

Design of a Pressurized Lunar Rover 24 April 1992 

S

Recorder 

SC Antenna 

• Analog data Subcarrier 

I oscillator 

Figure 9.2 - Telemetry Link 

L

9.4 Command Link. The command link (Figure 9.3) serves as the data up-link for 

the mission. It allows direct command from the base, with an option for command from 

earth. The data is transmitted digitally in a special command link code. The data is coded 

to prevent faulty messages or false commands. The signal is decoded aboard the rover, and 

• verified via telemetry back to the command link. Command reception triggers a command 

directly to the appropriate system. 

The decoder will require 17 W of power and the receiver will require 1.5 W. 

S 

I

103
S

Design of a Pressurized Lunar Rover 24 April 1992 

Onboard antennas 

S

Rec. I I Rec.

Combiner 

Sub-bit decoder . Verification signal 

to telemetry 

•

Message decoder I Message I distributor 

To subsystems 

Figure 9.3 - Digital Command Link 

9.5 Television Link. The rover's television link (Figure 9.4) is one of the less critical 

parts of the com'munications system. A slow scan of about 300 lines at 10 frames/sec will 

• suffice for most applications. By utilizing a digital scanning method, frame rate can be 

sacrificed for resolution (and vice versa), to better Suit specific applications. Figure 9.5 

shows picture quality levels for various signal-noise ratios (Sonnet). The ratios in the figure 

• show peak signal-noise ratios, not average ratios, which are about 3 dB less. It can be seen 

that 26 dB will be sufficient for good picture transmission, and the link will need about 12W 

of power. 

*

104 

5) 

L4 

11
._ ,; _4••

F

; 

Design of a Pressurized Lunar Rover 24 April 1992 

S-band antenna 

From camera I I Modulator I I Power 

driver amplifier 

Figure 9.4 - TV Link 

Figure 9.5 - Picture Quality Levels

Picture quality for different peak signal to rms noise ratio. 

9.6 EVA Link. Extravehicular activity will require special EVA communications 

links (Figure 9.6). This will be for short distances, around 100 to 1,000 meters. The EVA 

link will consist of a combination of a two-way voice link and biomedical telemetry. This 

can work on a voice-actuated switch; when the voice channel is not being used, telemetry 

105
e 

5 

El

S

Design of a Pressurized Lunar Rover 24 April 1992 

will be transmitted. This can be broadcast to the astronauts by an omnidirectional (0 dB gain) 

5) antenna on a VHF band, which will require minimal power. 

From sensors

Subcarrier I Telemetry 

oscillator '- subcarrier 

bank p...J combiner

Switch [---I Transmitter 

• From astronaut

IDuplexer 

El Receiver 

SUIT 

ONBOARD 

Figure 9.6 - Extra-vehicular Link

106
Design of a Pressurized Lunar Rover 24 April 1992 

9.7 Antennas. As previously mentioned, the EVA link will be transmitted by a small, 

low-power omnidirectional antenna, which will be mounted on the front half of the rover. 

All other communications links will be transmitted by a parabolic reflector dish. It has been 

shown that a 26 dB gain is adequate for picture, voice, and telemetry transmission. The gain 

of an antenna is related to its area by the following equation: 

G = 47riA/L2

where L = wavelength of transmission. 

With a 2.3 GHz signal broadcast by an antenna of 50% efficiency, the dish must have 

a diameter of 1.174 meters to produce a 26 dB gain. To account for antenna degradation, a 

1.2 meter parabolic reflector was selected, which will also be mounted on the front of the 

rover.

107
S

Design of a Pressurized Lunar Rover 24 April 1992 

S

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•

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•