Phobos Mission Vehicle Design

Department of Aerospace Engineering and Engineering Mechanics

The University of Texas at Austin

Guillaume Brat

Mustafa Khan

Jose Lozano

Walter L. Moguel

Kamran Moosavi

David North

Harold E. Pangilinan

December 10, 1991

Abstract

The preliminary design of an unmanned probe to the Mars moon Phobos has been developed. The primary goal of this mission is to gather chemical composition data with a particular emphasis on the search for water. Secondary goals are to completely map the surface of Phobos and investigate its internal structure via seismographic experiments. The mission will be accomplished with a three-module spacecraft including a propulsion module, a stationary lander, and a smaller mobile lander. Launch from Earth will be made with a Titan IV/Centaur G with a nominal launch date of July 9, 2009 and a total mission time of approximately one year. Preliminary design of vehicle subsystems includes structures, propulsion, thermal control, power, communication, guidance, navigation and control. In addition an analysis of trajectories has been made.

ObjectiveObjective

The purpose of the spacecraft is to explore the Mars satellite, Phobos. Surface landers will be used to gather data on material composition with particular emphasis placed on the search for water. The existence of water will justify additional missions to Phobos and will aid in future missions to Mars and more distant planets. Determination of Phobos' chemical composition may also contribute to the knowledge of the origin and evolution of the Solar system. Furthermore, the Phobos mission will act as a stepping stone for development of advanced planetary exploration techniques, which will aid in future manned explorations and habitation of Mars and Phobos

Mission and Spacecraft ProfileMission and Spacecraft Profile

Mission time from LEO to Phobos is about 260 days. During this period the RTG (on the stationary lander) is the sole source of power for the entire spacecraft. The penetrators are deployed during station-keeping about Phobos. These are designed to carry seismology and heat measurement sensors. The spacecraft (connected) scans for a good landing site for the mobile lander. Having found the landing site the mobile lander is deployed. The lander temporarily attaches itself to the surface (landing procedures are monitored closely to assure even more accurate landing procedures for the stationary lander). The lander remains at this first site for several days (perhaps a week). Meanwhile it constantly communicates with the stationary lander (still in station-keeping mode). At the end of this period the mobile lander departs from its original site and lands at another site (to see if the Phobos surface is similar within a reasonable range). The stationary lander separates from the propulsion module, lands, continues to track the first lander, and begins its experimentation. The propulsion module remains in station-keeping mode for the duration of the mission, inflating an echo ball with a concave surface to improve the local communication between the two landers when these are on the surface. The stationary lander runs its experimentation and monitors the environment for about three weeks. Total mission duration after having reached Phobos - about a month or 5 weeks.

Phobos EnvironmentPhobos Environment

Phobos was probed by Mariners 7 and 9 and Vikings 1 and 2. The surface morphology of the satellite has retained it's history of meteor strikes. The crater formation of depth and density has dated Phobos to be close to the age of the Solar System.

Phobos has an orbital radius of 9380 km with an eccentricity of almost unity. The shape of Phobos is a triaxial ellipsoid with semi-major axes of 13.4, 11.2, and 9.2 km. It has a gravitational constant m = 660 km3/s2. Surface temperature has been measured to be 305 K on the lit side and 190 K on the dark side. The surface radiation coefficient of Phobos is 0.9. Mean density is estimated to be 1900 kg/m3. It has been speculated that the carbonaceous chondrite material on Phobos may contain up to 20% water and 5% carbon.

Soviet Probes to PhobosSoviet Probes to Phobos

Two Phobos space probes were launched from the Baikonur Cosmodrone in July, 1988. The first craft was eliminated by erred telemetry. The second probe ran into technical problems but was able to establish an orbit about Mars just 100 km away from Phobos with a rendezvous period of about 8 minutes. The probe eventually encountered stabilization problems and the solar cells stopped supplying power. One of the most important outcomes of the Soviet mission to Phobos was the precise determination of its orbital elements.




Propulsion System and Trajectory DesignPropulsion System and Trajectory Design

The propulsion requirements for the Phobos mission can be broken down as listed below. The propulsion subsystems on the vehicle which will be used are indicated in brackets.

1) Vehicle launch and injection into the Earth-to-Mars trajectory [Titan IV/Centaur G]

2) Mid-course corrections and plane change in the Earth-to-Mars trajectory [Main Propulsion Engines]

3) Capture at Mars placing the vehicle in a near-Phobos "walking orbit" [Main Propulsion Engines]

4) Rendezvous (circularization) with Phobos [Main Propulsion Engines]

5) Stationkeeping at Phobos [Main Attitude Control/Station Keeping Thrusters]

6) Mobile lander descent and roving operations [Mobile Lander Propulsion]

7) Stationary lander descent [Stationary Lander Propulsion]

8) Attitude control during all of the above phases.

Trajectory Design/AnalysisTrajectory Design/Analysis

The Phobos spacecraft will be launched from Kennedy Space Center into a parking orbit 300 kilometers above the Earth. A short while later, a third-stage Centaur will ignite to boost up the speed by about 3.6 kilometers per second, flinging the 3800 kilogram spacecraft outwards towards Mars.

Two types of transfer orbits to Mars are considered for this mission. Type I is a broken Hohmann transfer with a simple plane change of 1.85 degrees at the point where the orbit plane of Mars crosses the ecliptic plane. This orbit is the one most economical in terms of both energy and flight time. The time of departure is set for July 9, 2005, and the mission will last for about 259 days. Upon arrival at Mars, the main engines will ignite to insert the craft into an equatorial circular orbit (9700 kilometers) around the planet. This orbit passes just 322 kilometers above its intended final orbit, which matches that of the planet's moon. Rendezvous with Phobos is accomplished by means of braking burns on opposite sides of the orbit. The total DV for the mission is 3.93 km/sec (not including launch vehicle/upper stage DV). Appendix A presents a detailed trajectory design and analysis of the mission.

The energy requirements of the mission can be reduced if aerocapture is used at Mars. In the "aerocapture" or "aerobraking" technique, a vehicle would plunge into a planet's upper atmosphere, where it would loose speed through air friction before emerging into the desired orbit. Preliminary analyses in Appendix A have revealed that energy gains as much as 1.20 km/sec can be accomplished if aerobraking techniques are used in the mission. However, due to increased complexity of techniques and systems and higher risk of mission failure, aerobraking is ruled out for this mission.

Type II transfer orbits use straight Lambert targeting to get to Mars. Lambert trajectories are not as energy efficient as Hohmann trajectories but they have the sole advantage of providing variant time-of-flights for a given mission.

For this mission, two constraints are placed on the trajectories. The first constraint limits the time-of-flight to approximately that of a Hohmann transfer (260-300 days). This constraint will assure a transfer orbit that is nearly tangent to the Mars' orbit upon arrival. The second constraint sets the arrival date at the first day of winter or the first day of summer. On these dates, the plane of Mars orbit coincides with the plane of Mars equator, therefore eliminating the need to equatorialize the orbit at capture.

Straight Lambert trajectory analyses are conducted for a 5-year period from year 2000 to 2005 with a step size of 30 days. The analyses show that orbits satisfying both constraints result in high DV's on the order of 40 km/sec. Substantial reduction in the value of these DV's can be accomplished if broken-Lambert trajectories are used. Low DV's on the order of 3.96 km/sec are obtained for orbits that only satisfy the second constraint. The time-of-flight for these orbits is 660 days. On the final analysis, a broken-Hohmann transfer is chosen as the best option for the mission.

Launch VehicleLaunch Vehicle

The launch of the Phobos vehicle will be made with a Titan IV/Centaur G launch vehicle. The 3800 kg fully loaded mass of the Phobos vehicle is well within the capability of the Titan IV/Centaur G for the chosen trajectory for this mission. A buffer of several hundred kilograms between the 3800 kg loaded mass of the Phobos vehicle and the total capability of the Titan IV/ Centaur G for the chosen trajectory/launch date has been allowed for the possible increased DV requirements of contingency trajectories/launch dates. The change in required DV between the chosen trajectory and contingency trajectories can be handled with a variation in the total impulse imparted by the liquid propellant Centaur upper stage.

Main Propulsion ModuleMain Propulsion Module

The four engines to be used on the main propulsion module are modified versions of the 870 lbf (N) thrusters used on the Space Shuttle. Maximum thrust for the main propulsion is 3480 lbf, however a maximum one g-load limit has been placed on the vehicle requiring an engine throttle back to approximately 25% of maximum thrust towards the end of the mission as vehicle mass is reduced. The engines are bipropellant engines using monomethyl hydrazine (MMH) as a fuel and nitrogen tetroxide (N2O4) as an oxidizer which are both space storable propellants. The propellants are hypergolic which eliminates the need for igniters. A thermal control system consisting of thermal blankets and solar energy collectors will be used to keep the propellants just above their freezing points (2o C) during the mission. Propellants will be fed from four main pressurized titanium propellant tanks pressurized to 1000 psi (6.895 MPa). Total Main Propulsion Module propellant tank volume is 1.86 m3 (65.5 ft3) [1.05 m3 MMH, 0.80 m3 N2O4], allowing 5% for ullage and unusable propellant. Total propellant mass is 2125 kg (4676 lbm)[ 669.5 kg MMH, 1456 kg N2O4]. The MMH tanks are 1.01 meters in diameter and the N2O4 are 0.914 meters in diameter. Total propellant tank dry mass (4 tanks) is 109 kg.

Pressurization for the main propellant tanks is provided by two smaller helium tanks which are 0.5 meters in diameter. The pressurization tanks also provide purge gas to the engines before and after firing. The helium pressurization gas is separated from the propellants in each of the propellant tanks by a flexible rubber bladder. Initial helium tank pressure will be 5000 psi (34.5 MPa) and will decay to about 1500 psi by the end of the mission.

Each engine will be cross-fed from all four propellant tanks to allow for complete propellant usage in the event of an engine failure. The main engine module can sustain three engine failures and still complete the mission. Additional redundancy is provided for control valves by placing parallel valves inline and for shutoff valves by placing additional valves in series. The addition of redundant valves and plumbing has a minimal impact on overall propulsion system mass (+8 kg). A schematic of the Main Propulsion Module is shown in Figure 1.








Figure 1. Schematic of the Main Propulsion Module

Main Propulsion Attitude Control/Thrust Vectoring/Station Keeping ThrustersMain Propulsion Attitude Control/Thrust Vectoring/Station Keeping Thrusters

Mounted on the outside corners of the main propulsion tanks are four attitude control pods each with four 20 lbf (89 N) thrusters for a total of 16 attitude control thrusters on the main propulsion module. The propellants for attitude control are the same as those for the main propulsion engines, MMH and N2O4. Propellants are fed from the main propulsion tanks. These thrusters provide attitude control around three mutually perpendicular axes of the spacecraft and main engine thrust vector control during the Earth-to-Phobos trajectory.

Station Keeping and Lander Deployment ManeuversStation Keeping and Lander Deployment Maneuvers

Upon rendezvous with Phobos the main propulsion station keeping thrusters will provide the DV needed to keep the vehicle positioned in an orbit around Mars approximately 2-3 kilometers above the surface of Phobos. Perturbations of this orbit by Phobos will require occasional burns to keep the vehicle within the 2-3 kilometer range. Propellant usage for station keeping at Phobos is estimated at 30 kg. At the times of lander deployment (landers are deployed at different times), the main propulsion unit will maneuver the landers to an altitude of 100-200 meters above the surface, deploy the landers, then move back to the 2-3 kilometer altitude. The landers will free-fall toward the surface after deployment. At an altitude of approximately 20 meters the landers will retrofire. Velocity at touchdown will be approximately 0.5 m/s.

Lander PropulsionLander Propulsion

The propulsion systems on each of the landers are identical. These systems act as both the attitude control system and descent engines. Both landers use hydrazine monopropellant blowdown systems. Pressurization for the hydrazine tanks is provided by separate helium tanks. Propellant tank pressure is held at 200 psi Each lander has sixteen 0.5 lbf (2.2N) thrusters providing 3-axis attitude control and translation. Estimated DV for descent and lander operations for each of the landers is 50 m/s which will require approximately 10 kg of N2H4 on each lander. The dry mass of each propulsion system on the landers is 8 kg.

TABLE 1

Propellant Usage

_____________________________________________________________

Event Isp DV Mo Mf DM

sec km/sec kg kg kg

_____________________________________________________

Centaur stage 450 3.6 23100 7300 15800

injection

Staging 7300 3800 3500

Trajectory

Corrections 280 0.03 3800 3752 48

Line Purge 3752 3742 10

1.85o Plane

Change 280 0.82 3752 3400 352

Line Purge 3400 3390 10

Trajectory 280 0.03 3390 3269 40

Corrections

Line Purge 3269 3259 10

Interplanetary 200 3259 3159 100

Attitude Control

Near-Phobos 280 2.7 3159 1181 1978

"walking orbit"

insertion

Line Purge 1181 1171 10

Rendezvous w/ 280 0.04 1171 1154 175

Phobos

Mobile Lander 200 0.05 200 190 10

Operations

Stationary 200 0.05 370 360 10

Lander Ops


TABLE 2

Propulsion Systems Mass Breakdown

Main Propulsion (not incl. support struct)

MMH tanks (2) 62 kg

N2O4 tanks (2) 47 kg

Engines (4) 60 kg

Piping 35 kg

Controller 10 kg

Valves 10 kg

Attitude control/vectoring thrusters(16) 8 kg

Helium pressurization/purge tanks(2) 36 kg

MMH 925 kg

N2O4 2011 kg

total dry mass 268 kg

total propellant mass 2936 kg

total mass 3204 kg

Mobile Lander Propulsion

N2H4 tank 3 kg

Piping 2 kg

Valves 1 kg

Helium pressurization tank 1 kg

Controller 1 kg

N2H4 10 kg

total dry mass 8 kg

total.propellant mass 10 kg

total mass 18 kg

Stationary Lander Propulsion

N2H4 tank 3 kg

Piping 2 kg

Valves 1 kg

Helium pressurization tank 1 kg

Controller 1 kg

N2H4 10 kg

total dry mass 8 kg

total.propellant mass 10 kg

total mass 18 kg

__________________________________________________

total propulsion dry mass 284 kg

total propulsion propellant mass 2956. kg

total propulsion mass 3240. kg

Power SubsystemPower Subsystem

The power system design for the Phobos mission has been sized for the stationary and mobile landers at a limit load of 200 W for each lander. The stationary lander will implement a modular type Radioisotope Thermal Generator (RTG). The mobile lander will use a solar array/battery combination power design.

On the stationary lander, a current of 28 V will distribute a power of 200 W produced by a 10 module RTG at a weight of 22 kg as shown in the figure below. This RTG will be in continuous operation from launch to end of mission on Phobos. The lifetime for such an RTG is 8 years. This system will also be the sole power source enroute to Phobos from earth to save the use of the solar arrays on the mobile lander upon landing at Phobos to minimize degradation effects. The stationary lander will continue to use the RTG after landing on Phobos. (See Appendix B for RTG design.)


Figure 2. MITG (Modular Isotopic Thermoelectric Generator)

The mobile lander will use Gallium Arsenide (GaAs) cells which is the latest improvement in solar cell technology. The GaAs cell provides an improvement of 18% efficiency over the standard silicon cell of 11.5%. This allows for a reduced area size of the solar array. The array will be deployed in a flexible fold-out light-weight configuration. The total area of the solar array is 3.4 m2 which produces a power load of 200 W plus the power necessary to recharge the batteries. The mobile lander will use 4 solar panels, each of dimensions 0.5 m x 1.7 m. The solar panels will pivot along the longitudinal axis, using sun-angle sensors to maximize the sun's solar intensity. (See Appendix B for Solar Array design.)

For the 4 hour nights on Phobos, Nickel Hydrogen rechargeable batteries will provide the necessary power load to the pertinent systems within the mobile vehicle. During the day time, the batteries will be recharged by the excess power provided by the solar panels. A fully regulated power bus will be used to control the distribution of power during recharge and discharge of batteries. A shunt regulator will regulate the distribution of power to recharge to maximize the life of the batteries. A discharge regulator will regulate the battery voltage discharge. (See Appendix B for Battery design.) Mass and sizes of the power system are shown in Table 3.


Table 3. Power Subsystem Mass and Sizes

Stationary lander-

Module RTG 22 kg 20"x18"

Mobile lander-

Solar Array 4.2 kg 2-3'x6'

NiH2 Btry 20 kg

Regulated Bus 4.3 kg

Guidance, Navigation and ControlGuidance, Navigation and Control

From earth to final Phobos rendezvous guidance and navigation will be controlled by a sun sensor (0.2 kg, 0.2 W), star tracker (10 kg, 5 W) and ring laser gyros. Control will be provided by four attitude control pods positioned on the propulsion module. Once at Phobos, each lander will also require autonomous guidance and control as many of the landing and surface activities will preclude constant communication with earth. Control on each lander will be accomplished with the use of control thrusters (as previously discussed in the propulsion section of this report).

Computer SubsystemComputer Subsystem

The stationary lander computer system is the core of the spacecraft computer subsystem. Once on Phobos, it is used not only as a relay for communication with Earth, but also to choose an appropriate landing site, and eventually to palliate any deficiency of the mobile lander. The computer on the stationary lander will serve as the central command computer during most of the mission. The mobile lander has just to handle data acquisition and storage, some communications with the stationary lander, and hopping commands in the last phase of its mission. Therefore, the mobile lander makes use of a simple computer system which requires low power consumption and low weight. In fact, it is only able to handle basic tasks, and will let the stationary lander computer system handle the heavy work. Moreover, the first landing may fail due to an underestimated parameter such as Phobos gravity, or surface density thus further justifying the use of a simple computer for the mobile lander.

Mission RequirementsMission Requirements

In the early stage of cruise to Phobos, the stationary lander computer system will be turned on. During cruise it will be in charge of communicating with Earth for trajectory correction purposes. As no mission experiments will be conducted before the arrival in Phobos, most of the memory will be available to store parameters needed for control of the vehicle during cruise. Although these parameters may be predefined, we do not use the ROMs (Read Only Memory) to store them since they can be subject to modifications. Indeed, modifying a ROM is quite expensive in terms of time and power. The ROM memory is divided into three parts. The first one contains the routines that handle any operations concerning communication with Earth. The second part contains the routines necessary to control the environment (temperature, power, etc...) of the spacecraft. The last one stores all the routines handling trajectory and attitude control.

When the spacecraft arrives at Phobos, the part containing the routines for controlling trajectory is erased and replaced by routines needed to control the landing of the mobile lander as well as routines needed by the stationary lander for communicating with the mobile lander. As the communication window between Earth and Phobos is very small (17 min), some of the routines will be loaded during the cruise stage just prior to the arrival at Phobos. Once in orbit, trajectory routines will be erased and replaced by the rest of the routines. At this time, the computer mobile lander system is turned on so that the landing can be controlled by the lander itself. The ROM memory of the mobile lander is loaded from the beginning with all routines necessary for its missions.

The next phase is characterized by the beginning of scientific experiments by the mobile lander. The mobile lander computer system is concerned mainly by collecting the data acquired from the experiments, compacting them and sending them to the stationary lander before each communication window. It also controls the environment of the mobile lander. Meanwhile, the stationary lander computer is in charge of communicating with Earth the scientific data (when it is time) and controlling its own environment as well as its attitude.

Once the mobile lander has conducted most of its experiments, the stationary lander has to chose a landing site. This site is decided on Earth after analysis of some of the data already received from the mobile lander. The appropriate landing coordinates are then sent to the stationary lander. While the stationary lander is landing, all communications from the lander are ignored. The link is reestablished once the landing is successful. As in the previous phase, compacted data stored in the mobile lander are sent to the stationary lander that transmits them to Earth with its own data. At this time, the mobile lander has carried on most of its assignment and is storing less data.

In the last phase, data acquired by both the stationary lander and the mobile lander will require less and less computation time of the stationary lander computer system. Therefore, the stationary lander computer system will concentrate on monitoring the moves of the lander. Both computers systems will still be controlling their environment and handling their communications.

Computer TechnologyComputer Technology

The computer subsystem of this spacecraft uses the GVSC (Generic VHSIC Spaceborne Computer) computer of Honeywell. The GVSC is based on a custom processor (Honeywell 1750A). In order to keep a low workload, three CPUs and two FPPs cooperate for executing an instruction. One CPU is in charge of the instruction pipeline, another one prepares memory addresses, and the last one deals with ALU functions. One FPP is dedicated to additions and subtractions while the other one handles multiplications and divisions. Our custom GVSC provides 8 MBits of RAM memory protected by EDC circuit. It uses RICMOS (Rad-hardened Inverted CMOS). It also offers 512 KBits of modifiable ROM (EEPROM).

The card cage has very small dimensions (8" x 7" x 4.5") because the GSVC uses VHSIC technology which reduces overall board size and power consumption. It can tolerate radiation (MIL-STD-1750) up to 1E6 rad (Si). The use of VHSIC technology guarantees good mechanical resistance. The power consumption is about 10 Watts for the CPU card with memory. The weight is 6 kg. The operating temperature should be kept close to 27oC. However, it can still operate efficiently in the range -10oC to +50oC.

Stationary Lander Computer SystemStationary Lander Computer System

As we explained previously, the stationary lander computer system is the most important part of the spacecraft computer subsystem. Therefore, it is composed of three processor units (i.e., three GVSC processor boards). They are arranged in a duplication with a spare redundancy technique. The duplication is realized by two processor units organized in a Master-Slave architecture: they both compute the same results; the Master compares both results and decides if it should declare itself down or if they should go on with the same configuration. In the presence of faults, computations are duplicated or even triplicated (if time is not an issue) so that transient faults can be detected and eventually corrected. Reconfiguration occurs only when a permanent fault is detected. In this case, the spare replaces the failed unit and is playing the role of the Slave.

The total weight of this system is 18 kg. The total power consumption is less than 30 watts. The dimensions are 8" x 7" x 4.5".

Mobile Lander Computer SystemMobile Lander Computer System

The mobile lander computer system is not very different. It is composed of two processor units arranged in a Master-Slave configuration. Like the stationary lander computer system, they use retry techniques to detect and correct transient faults. Upon permanent fault detection, the mobile lander computer system asks for arbitration from the stationary lander computer system. This one computes the last computation done by the mobile lander computer system and decides which unit should be disabled. After the first permanent failure, the lander computer system functions as a single processor unit system. It relies only on built-in schemes to detect failures.

The total weight of this system is 12 kg. The total power consumption is less than 20 watts. The dimensions are 8" x 7" x 4.5".



Figure 3. Stationary lander computer system interface.







Figure 4. Mobile lander computer system interface.



Communication Subsystem

Like the computer subsystem, the communication subsystem is distributed on the stationary lander and the mobile lander. The stationary lander communication system is the most important one since it takes care of all communication with Earth. The mobile lander relies entirely on the stationary lander to communicate with Earth. Therefore, both modules have an identical local communication system based on an omni-directional low-gain antenna. The stationary lander has also a high-gain antenna (similar to Viking's antenna) for communicating with Earth. While remaining in stationkeeping mode, the main propulsion module of the spacecraft deploys an echo ball with a concave surface to improve the local communication between the two modules when these are on the surface.

Local Communication SystemLocal Communication System

Communications between the stationary lander and the mobile lander are transmitted via low-gain omni-directional antennas. As they will never be very far from each other or from the echo ball orbiting around Phobos, the chosen radio frequency band is UHF. The amount of communications is not very high since data are first stored in the mobile lander memory and then send to the stationary lander. Usual transponders are used to receive and transmit the signal. They provide an output power of less than 1 watt.

Each low-gain communication system consumes about 40 watts. The total mass is estimated at less than 10 kg: the antenna itself weights less than 3 kg. The size of the antenna is about 0.7 m of diameter. A data rate of 1200 bauds should be enough to transmit experiments data and eventual commands.

Earth Communication SystemEarth Communication System

The long range communication system is similar to the one used for the Viking mission. The high-gain antenna operating in S-band or X-band frequency is able to accommodate a higher data rate than the local communication system. This is necessary since the communication window lasts only 17 minutes. In this lapse of time, the stationary lander should be able to transmit the data stored in memory and receive eventual commands from Earth. Therefore, a two-way antenna and different frequencies for receiving and transmitting a signal are used. This system uses NASA standard transponders which produce an output power of 20 watts.

The size of the antenna is about 1.7 m in diameter. The total weight of the long distance communication system is about 100 kg. It consumes 40 watts of power. The uplink frequency is 2680 MHz in S-band and 8100 MHz in X-band. The downlink frequency is 2520 MHz in S-band and 7500 Mhz in X-band. A data rate of 9600 bauds allows the stationary lander to transmit 9,792,000 bits during 17 min. This means than about 1 MByte of data can be sent each day to Earth.



Figure 5. Communication when the stationary lander is in orbit.










Figure 6. Communication when the stationary lander has landed.



Scientific SensorsScientific Sensors

Sensor GoalsSensor Goals

The scientific sensors on board this mission will attempt to answer some important questions about Phobos. The surface and subsurface composition at this time can only be inferred from meteor impacts, orbit, volume, mass, and luminosity. Scientists can only speculate the composition of the core. The heat flow and seismic data is nonexistent. This mission may bring the discovery of water bearing rock. Accessible water on Phobos can lead to an early colonization of Mars or at least a permanent scientific presence on Mars.

Sensor ManifestSensor Manifest

The scientific sensors to be carried are modeled after the Surveyor and Luna probes to the surface of the Moon. The present instruments were carried on the Phobos probes launched in 1988. The neutronmetry, ion mass spectrometer, X-ray mass spectrometer, and the electric field meter are operated consecutively not concurrently. The rest of the sensors list below are operated concurrently. The landers will carry,


Stationary Lander Scientific Sensor Manifest

Neutronmetry, subsurface mass spectrometry 10 W 18 kg

Ion mass spectrometry for soil surface 5 W 5 kg

X-ray radiometry for soil surface 10 W 10 kg

Heat Flow Experiment 5 W 6 kg

Magnetometer 5 W 5 kg

Solar Wind Spectrometer 5 W 5 kg

Electric Field Meter 10 W 20 kg

Visual Imagery 15 W 20 kg

Three Axis Seismometer 5 W 5 kg

Stationary lander total scientific sensor mass 94 kg

Stationary lander scientific sensor

continuous power consumption 40 W

Roving Lander Scientific Sensor Manifest

Laser mass spectrometry for soil surface 10 W 70 kg

Heat Flow Experiment 5 W 6 kg

Magnetometer 5 W 5 kg

Three Axis Seismometer 5 W 5 kg

Roving lander total scientific mass 86 kg

Roving lander scientific sensor

continuous power consumption 15 W

PenetratorsPenetrators

The probe will carry six penetrators. Each penetrator is equipped with heat flow sensors employing thermistor arrays, a communication system, spin booms, solid motors, and spin thrusters. Three penetrators will have an explosive charge, minimum communications, minimum battery power, heat flow sensors, and nose cone ballast. The three other penetrators will be equipped with a three axis seismometer with a sensitivity of 1 Hz, long term batteries, communications, data storage, and heat flow instruments. In these three penetrators the batteries take the place of the charges and the seismometer instrumentation replaces the ballast load in the nose cone.

Heat Flow MeasurementsHeat Flow Measurements

The heat flow measurements are accomplished with the use of two arrays of 12 thermistors circumferentially lining the penetrator hull. One array will be mounted just aft of the nose cone. The second array is mounted aft of the radio transmitter. Both arrays are insulated from internal heat sources. In addition, internal heat sources such as communications, batteries, and certain parts of the seismometers are also insulated to prevent interference with the thermistors.

Some of the instrumentation will be hardened against high deceleration by submersion in a gel or epoxy depending on existing technology. The electronics systems do not need to be shielded from radiation since the depth of penetration should be a sufficient deterrent to the effects of radiation.

The spin booms will act as antennae for the communications system. At impact, the spin booms will detach from the penetrator hull and remain on the surface or at least near the surface. Cable wiring are similar to the trail wiring found on anti tank TOW missiles developed by Hughes Inc. The spin booms are deployed by clock springs when the penetrator is released by the lander. If the springs do not deploy the booms, the inertia of the spin booms during the propellant burn will enable its deployment. Two valves located on the motor nozzle and directed in a tangential direction instigate the spinning motion of the penetrator. The spin booms stabilize the spin and maintain the spinning motion during flight.

The solid propellant motor on the penetrator will be able to produce a DV of least 150 m/s. The penetrator hull will be constructed from titanium. Hull design allows for proper penetration and a reduced deceleration environment. The hull design accounts for a wider angle of incidence range that would prevent the penetrator from skipping off the surface of Phobos while allowing for volume and mass constraints.

The penetrator will be ejected by the probe using a spring mechanism. The spring mechanism will initiate the spin for the penetrator. The spring induced spin will stabilize the penetrator until it reaches a safe distance from the orbiter where it can ignite it's solid motor. The penetrators will be mounted tube style in a six tube cluster.

The penetrators will be fired after the probe has entered the rendezvous orbit with Phobos. All six penetrators will be deployed before the main lander deploys. The three penetrators that serve as seismometer stations will be powered by lithium batteries. Expected lifetimes for the seismometer equipped penetrators will be between 6 months to a year. They will relay data to the main probe using the spin booms as omni directional antennas. The spin booms detach from the penetrator at impact. The spin booms remain on or near the surface. The spin booms are connected to the radio transmitter through wiring which was unspooled during impact. Seismometers will send a data rate to the communications system at 2 kbits/sec. The communications system will send data in bursts storing excess data in a one megabyte DRAM chip. The probe will continuously gather data until the data storage has reached it's capacity. Data acquisition resumes after the data transmission burst to the orbiter. Variation of the time lag between the end of the transmission burst and the beginning of the data acquisition routine, can be adjusted through instructions sent during the transmission burst.

Despite the present antenna design, the three radio activated penetrators containing charges may encounter radio reception difficulties. To counter this problem, the penetrators will be detonated before the orbiting probe deploys. Should the penetrator not receive detonation instructions, a timer delayed set of instructions programmed into the penetrator will serve as a backup. Information from the artificially generated seismic waves are picked up by the first lander and the three seismometer equipped penetrators. The seismometers will continue gathering data from thermally and gravitationally induced seismic activity.

Penetrator mass without motor

and spin booms 13kg

Penetrator total mass 25kg

Total weight of six penetrators 150kg

Laser Mass SpectrometerLaser Mass Spectrometer

The system emits a focused laser at the surface of Phobos which will cause the soil to vaporize. The vaporized cloud emits ions which are picked up by a reflectron on the craft by measuring the time of flight of the ions. The reflectron has a retarding electric field which slows the displaced ions for time measurements. The spectrometer will be designed to determine the elemental and isotopic composition of the surface. It is capable of measuring, up to 200 atomic mass units. 200 mass units is approximately in the heavy metal range on the table of elements. The resolution of the sensor is rated with respect to the masses that are being measured. So an absolute measurement of it's resolution cannot be determined beforehand. The laser used in this system has a wavelength in the infrared region. Specifications for the sensor are as follows,

Range of masses measured 1-200 amu

Resolution with respect to measured masses 150 M/DM

Wavelength of the laser 1.06 mm

Energy per pulse .5J

Duration of pulse 10ns

Frequency of pulses .1-.2 Hz

Mass of the sensor 70kg

The laser mass spectrometer will have a range of 50 meters. With this range, the spectrometer will be mounted on the roving lander. It will be mounted on the top surface of the roving lander. The turret mount of the laser mass spectrometer will have a 360 degree rotational freedom. The laser mass spectrometer will have a preprogrammed firing and measuring sequence of the environment. The sequence is executed whenever the lander has acquired a new position. Instructions for further firing sequences for better resolution, can be executed based on the data gathered from the preprogrammed sequence.

Ion Mass SpectrometryIon Mass Spectrometry

The sensor will determine the elemental and isotopic composition of the soil. A krypton ion beam fires at the soil which bumps out ions in a cloud form present within the soil. The displaced ions are analyzed by a mass analyzer. The instrument will yield information about elemental and isotopic composition of the surface. The instrument may yield data on the history that has taken place on the surface. It is capable of measuring up to 60 atomic mass units which is the atomic mass of Neodymium. With this range it is capable of detecting all the common rock elements except the heavy metals. Its resolution is also relative to the measured masses. Specifications of the sensor are listed below,

Range of ion masses recorded 1-60 amu

Pulse duration 1 sec.

Repetition frequency .2Hz

Beam current 2 mA

Energy of the ion beam 2-3keV

Mass of the sensor 18kg

The ion mass spectrometer will be mounted on a revolving turret on the bottom surface of the stationary lander. It will be able to analyze the composition of the surface directly below the stationary lander. The sensor will have a preprogrammed firing and detection sequence which will be executed once the lander has entered a range of 50 meters from the surface. There will be another sequence of detection firings once the probe has landed. After the sensor has performed it's duty, the sensor will be turned off.

NeutronmetryNeutronmetry

The neutronmetry requires the use of a neutron flux generator and a scintillation gamma ray spectrometer. The neutron generator fires a flux of 14MeV neutrons with pulses ranging from 1 to 3 msec at a frequency of 10 kHz. The striking neutrons cause the rock to reradiate which is picked up by the scintillation gamma ray spectrometer.

The neutron pulse generator consists of a cylindrical electrovacuum tube. Neutrons are generated by accelerating deuterium ions up to 100 keV while absorbing tritium atoms on zirconium target. The pulse generator will carry a large array transformer which will generate up to 100 kV. The transformer requires similar technology and weight characteristics as an inductor coil found on most GM distributorless ignition systems for automotive engines.

The detection unit will double as a scintillation unit that will pick up the cosmic ray and natural gamma ray radiation and the detection unit that will pick the gamma ray radiation induced by the neutron pulse generator. The transformer unit for the detector will boost the voltage to 1800 V. To protect the detection unit, the transformer units for the neutron pulse generator and detector will serve as a shield against the neutron pulse generator.

Measurements from the neutronmetry system will provide mass spectroscopy of the soil up to 20 meters in depth depending upon the local density of the soil. Detection resolution decreases with increasing depth. Rates of resolution degradation vary as a function of the atomic mass of the elements detected.

The neutronmetry system will be mounted on the bottom mounted swivel turret on the stationary lander. Then a preprogrammed firing sequence will be executed after the all the mass spectrometers have taken data. Each neutron pulse firing and detection cycle lasts about 45 minutes. A preprogrammed firing sequence is executed after the ion mass spectrometer and X-ray radiometry experiments have been completed.

X-Ray RadiometryX-Ray Radiometry

The X-ray radiometer radiates the rock with high energy X-ray. The X-ray excites the rock atoms and ionizes the atoms. The ionization occurs at certain energy levels characteristic for each element. The repetition of each energy level is proportional to the concentration of each element.

The X-ray radiometer is mounted on the bottom surface of the lander away from the swivel turret. Each spectrometer on the lander will operate consecutively not concurrently.

Other SensorsOther Sensors

Measurements of the soil temperature will be accomplished using thermistors. Three array thermistors are mounted on all the landing pods of the stationary and roving landers in order to facilitate a solid contact with the surface. Magnetic measurements are executed by ferroprobe magnetometers which measure the change in phase of magnetic waves. Power consumption of the sensor will be in the range of 30 to 65 mA. The magnetometer will be operated after all the mass spectrometers have completed their tasks. Also, the magnetometers will operate on both landers prior to landing. An optical image sensor mounted on the stationary lander, will be used for transmission of images of the surface of Phobos. In addition, seismometers will measure the seismic activity induced by the exploding penetrators. The stationary lander, mobile lander, along with penetrator mounted seismometers will determine the density of the core. The seismometer will measure the thermal expansions and gravitational stretching and compression due to the Martian gravitational field.












Structure and ConfigurationStructure and Configuration

The spacecraft is comprised of three main structures: the propulsion module, stationary lander, and mobile lander. The maximum width and length of the spacecraft is 3.0 m and 5.5 m respectively. The mobile lander is stacked on top of the stationary lander, which in turn is positioned above the propulsion module. For the launch configuration, the propulsion module forms the base of the entire spacecraft assembly. The landers are positioned right side up (i.e. landing gears pointing down towards the propulsion module). The interfacing between the three main structures is provided by short trusses of boron/epoxy tubes.

The interfacing trusses are characterized by short, vertical support struts to add rigidity to the spacecraft. The truss connecting the mobile lander to the stationary lander is just long enough to safely accommodate any equipment (i.e. cameras, a support structure for the high gain antenna, and GN&C equipment) that would be positioned on the top deck of the stationary lander. The truss connecting the stationary lander to the propulsion module is also characterized by short, vertical support struts. The propulsion module is a truss of boron/epoxy tubes and struts structured around the four 1.0 m wide spherical propellant tanks.

All support struts and tubing are primarily designed against buckling loads and are fabricated from a boron/epoxy composite. Boron epoxy is a composite material "with unidirectional fibers of high elastic modulus [that] are particularly advantageous in saving weight", (Meyer, U.C.L.A.). The main structural bodies of both landers are constructed primarily from machined aluminum and titanium, a material combination that is modeled after the Viking lander body structure (Holmberg, 1980).

Detailed drawings of the entire spacecraft and of each lander are provided in the appendix.

Propulsion ModulePropulsion Module

The propulsion module is composed of the following:

1. Eight 1.353 m sections of tubing (see diagram)

2. Two 1.9 m sections of tubing for the top and bottom truss sections

3. Two 2.234 m sections of tubing (cross cube diagonals)

4. Four 0.33 m sections of tubing (see diagram)

5. Four spherical tanks (109 kg)

6. Four thruster assemblies (60 kg)

Total mass (approximated): 51.5 kg (tubing) + 109 kg + 60 kg = 220.5 kg.

Given the lengths of boron/epoxy tubing with an approximate 4.0 cm diameter (density: 2.01 kg/m3 x 1000), and the masses of the spherical tanks and thruster assemblies, the mass of the propulsion module is approximately 220.5 kg.


Figure 14. Propulsion Module

Stationary LanderStationary Lander

The stationary lander is a two level structure. The bottom level is a six-sided box, 0.66 m in depth, with alternating 1.155 m and 0.165 m sides. The top level is triangular with a depth of 0.33 m and 1.0 m sides. The top triangle is offset from the bottom six-sided box to allow for placement of cameras and GN&C equipment (including the needed Sun and star trackers). The body is constructed primarily of aluminum (density: 2.8 kg/m3 x 1000). Its dry weight (not including equipment, propellant structures, and landing legs) is 30.0 kg assuming wall panel thicknesses of 2.0 mm. The mass of the landing legs is approximately 11.43 kg. The lander houses:

1. Spherical tank of N2H4 (hydrazine) of 0.25 m diameter and a smaller tank of pressurized Helium. Both tanks are made from titanium.

2. Three 6 kg computers

3. One 86.6 kg Nickel-Cadmium battery

4. Two swivel mounted turrets containing the following data acquisition equipment:

Bottom turret:

Neutronmeter

Ion mass spectrometer

X-ray radiometer

Top turret:

gamma ray spectrometer

Solar Wind spectrometer

Electric Field Meter

The two swivel mounted turrets are housed inside the lander body (one on the bottom level and the other on the top level). When needed for data acquisition, panels on the bottom and top level boxes will open, thus allowing the turrets to safely move into position (see diagram below). By enclosing the sensor turrets the sensitive equipment is protected from possible damage during the landing procedure (the composition of the Phobos surface is relatively unknown, thus the consequences of a 0.5-1.0 m/s landing is also unknown).


Figure 15. Sensor Turrets

The stationary lander stands 0.66 m above the surface when the landing legs are fully deployed. Due to the low descent velocities of the lander, the landing legs require a minimum degree of impact attenuation properties (which eliminates the need for large high-pressure hydraulic struts). The landing pads are approximately 0.4 m in diameter. The wide footprint of the individual landing pads is intended to model the wide footprint provided by snow shoes. This would effectively prevent the lander from sinking into the surface (if this could indeed happen). Each landing pad is also equipped with small surface anchoring penetrators. These are deployed in order to secure the lander in place.

The stationary lander body also provides mounting provisions for the boom mounted RTG (22 kg), penetrator launcher assembly (150 kg), high gain antenna, two cameras, and four attitude control thrusters. The boom for the RTG is constructed of steel and titanium. The choice of materials for the RTG boom is based on design specifications from past deep space probe missions, such as Voyager, which also made use of RTGs as sources of power (NASA, 1977). During the launch configuration, the RTG boom is rigidly attached to the payload fairing of the launch vehicle (NASA, 1977). The boom is deployed immediately after the spacecraft's separation from the launch vehicle in order to effectively protect the sensitive onboard equipment from the intense radiative and thermal properties of the 200 Watt RTG. The RTG's dimensions (including heat radiating fins), are 18.0 in. (0.46 m) x 20 in. (0.508 m). Mounts for the two cameras are machined into the lander body structure. The attitude control thrusters are rigidly attached to the structure.

Mobile LanderMobile Lander

The mobile lander body is approximately 0.457 m or 1.5 ft in depth. Its shape is octagonal with 0.457 m sides. Like the stationary lander, the body is constructed primarily of aluminum. Its dry weight (not including equipment, propellant structures, and landing legs) is 23.55 kg. The mass of the landing legs is approximately 11.43 kg. The lander houses:

1. Spherical tank of N2H4 (hydrazine) of 0.25 m diameter and a smaller tank of pressurized helium. Both tanks are made from titanium.

2. two 6 kg computers; and

3. one swivel mounted turret containing the laser mass spectrometer. When deployed the turret is found on the top side of the lander (deployment of this turret is similar to the deployment of similar turrets on the stationary lander).

The mobile lander stands 0.66 m above the surface when the landing legs are fully deployed. The landing legs geometries are identical to those on the stationary lander but with an added feature for mobility. Each landing pad is equipped with four small surface anchoring penetrators. While the stationary lander requires one set of anchoring penetrators to serve as permanent anchors, the mobile lander's need for stability during data acquisition and mobility for roving requires several sets of penetrators. The penetrators will be deployed for securing the lander to the surface. For mobility, the lander will release itself from the penetrators. Upon landing at another site, the lander will deploy the next set of penetrators.

The mobile lander body provides mounting provisions for four 0.5 m x 1.7 m gallium arsenide solar panels (for a total solar panel surface area 3.4 m2) and four attitude control thrusters (distributed symmetrically along the circumference of the lander body), two cameras, and a low gain antenna. The solar panels are deployed during surface operations on Phobos and stowed when the lander becomes mobile.

Thermal SubsystemThermal Subsystem

The temperatures within all the subsystems will be monitored throughout the mission. The temperature control of the design will maintain all parts of the spacecraft within operating temperature limits for the range of conditions from launch to extended stay at Phobos.

The design of the thermal subsystem will employ both active and passive techniques to create the necessary environment .

Active:

(1) Bimetallic actuated louvers

(2) Solar energy collectors (SECs)

(3) Gas pipes from RTGs

Passive:

(1) Paints and coatings

(2) Multilayer insulation blankets (MLI)

(3) Thermal capacitance and conductance of structure

A table showing operating temperature limits of the subsystems and equipment is given below:

Table 4. Typical equipment temperature limits
Thermal Design
Temperature
Limits (°C),
Min/Max
Subsystem/Equipment
Nonoperating/Turn-on
Operating
Communications

Receiver

Input multiplex

Output multiplex

TWTA

Antenna


-30/+55

-30/+55

-30/+55

-30/+55

-170/+90

+10/+45

-10/+30

-10/+30

-10/+30

-170/+90
Electric Power

Solar array wing

Battery

Shunt assembly


-160/+80

-10/+25

-45/+65

-160/+80

0/+25

-45/+65
Altitude control

Earth/sun, sensor

Angular rate assembly


-30/+55

-30/+55

-30/+50

+1/+55
Momentum Wheel
-15/+55
+1/+45
Propulsion

Solid apogee, motor

Propellant tank

Thruster catalyst bed


+5/+35

+10/+50

+10/+120

-

+10/+50

+10/+120
Structure

Pyrotechnic mechanism

Separation clamp


-170/+55

-40/+40

-115/+55

-15/+40

It is seen that an average equilibrium temperature of 20° C (293° K) is suitable to maintain the spacecraft within these limits.

The biggest problem encountered by the spacecraft from a thermal point of view is the large difference between solar flux radiation at Earth (1362 W/m2) and at Mars (590 W/m2). This problem will be overcome by shielding the spacecraft from solar radiation. This will be accomplished with the use of insulation blankets, paints and coatings.

The spacecraft can be divided into four areas,

(1) Propulsion Module

(2) Stationary Lander

(3) Mobile Lander

(4) Appendages.

Propulsion Module Thermal ControlPropulsion Module Thermal Control

The propellants and pressure tanks need to be maintained at 20°C and cannot go below 0°C. For temperature control the tank assembly and pressure tanks will be enclosed in multilayer insulation (MLI). The temperature of the tanks will then be controlled by four solar energy collectors (SEC).

The MLI consists of several layers of closely spaced radiation reflecting shields which are normal to the direction of heat flow. The aim of the shields is to reflect a large percentage of the radiation the layer receives from warmer surfaces thereby insulating the spacecraft.

The insulation blanket covering the propulsion module consists of an extremely thin outer layer made of aluminized Kapton (absorptivity, a = 0.35 , emissivity, e = 0.6) and inner layers made of aluminized Mylar. Dacron mesh is used to separate each of the layers.















Figure 16. Multilayer Blanket Composition

The surface of the MLI is covered in aluminum foil (a = 0.2, e = 0.04) except for the underside of the blanketing case where the thrusters are located.




Figure 17. SEC Positioning on Exterior of MLI



The SECs will be positioned on the exterior of the MLI (Fig. 17). The SECs work like louvers and can vary the amount of solar radiation intake by opening or closing the covers. The graph below shows the net energy captured by each SEC for several solar intensities between Earth and Mars as a function of opening angles.


























Figure 18. Energy vs. SEC Opening Angle

The main propulsion thrusters will be painted black (a = 0.93, e = 0.88) and the expected temperatures of the tanks are:

At Earth (°C) At Mars (°C)

Aft tank shell 25 20

Forward tank shell 21 17


Stationary Lander Thermal ControlStationary Lander Thermal Control

The temperature of the stationary lander will be controlled using a gas pipe coming from the RTG and bimetallic actuated louvers. Three sets of louvers will be spaced 120 deg around the lander, and each painted with white Silicone paint (a = 0.14, e = 0.88) on the exterior and interior. When the louvers are closed the white exterior will reflect the solar radiation while the white interior will act as an insulator. The rest of the lander will be painted with Aluminum Silicone paint (a = 0.25, e = 0.28).

The heat required to maintain the desired temperature will be provided by a pipe of gas from the RTG. The RTG will be on a deployable boom. Its heat output will be about 1800 W. The gas will be heated using this heat and then pumped to the lander. The pipe will go around the base plate of the lander and will provide heat to the instruments inside. Once the pipe has transferred the heat the cooled gas will return to the RTG for reheating and recirculation. The pipe itself will be completely wrapped in an insulation blanket except the portion that goes inside the lander which will be left uncovered

The fluid chosen is a gas and not a liquid because if for some reason the temperature fell sharply the liquid could freeze and heat to the lander would be cut off. However gases have much lower freezing points than liquids and thus overcome this problem. The choice of gas is dependent on two factors. Firstly its chemical properties and secondly its thermal conductivity. To prevent any possible chemical reactions between the gas and the inner surface of the pipe an inert gas will be used. Gases also tend to have low thermal conductivities so any inert gas would be suitable. Argon was chosen (k = 0.016 W/mK) since it is relatively more available.

All the instruments inside the lander will be individually insulated and wrapped in blanket insulation. The Hydrazine tanks will be insulated with MLI. The cameras are located on top of the lander and so in addition to insulation also have replacement heaters. When the cameras are operating the replacement heaters will be off since the heat given off by the cameras is retained by the insulation. However when the cameras are off the replacement heaters will come on and provide 1.75 W of heat to each camera.

When the temperature inside the lander starts rising above 25°C the louvers will be opened using a bimetallic actuated spring. When the temperature starts falling below 5°C the louvers will close. The temperature of the stationary lander will be controlled in this way during cruise and after landing on Phobos since the RTG boom will be attached to the stationary lander.

Figure 19.

Mobile Lander Thermal ControlMobile Lander Thermal Control

During the cruise portion of the mission the mobile lander will have the same temperature control system as the stationary lander, i.e heated gas pipe from the RTG and bimetallic actuated louvers. On the mobile lander four sets of louvers will be spaced at 90 deg around the lander. The paints and coatings of the exterior will be exactly the same as for the stationary lander and all the instruments will be individually insulated.

The difference between the two landers occurs after separation, from which point on the deployed solar panels will provide power to an electrical heater inside the lander. This heater in conjunction with the louvers will then provide the temperature control of the lander.

It is found that despite the slight mass penalty it is better to have two separate pipe systems for the landers. If there was only one pipe circuit around the whole spacecraft severe problems would be encountered on separation prior to landing. There would be several break off points which would need to be sealed and then a new pipe circuit would have to be made for the stationary lander. To avoid these complications it is better to have two separate pipe systems. On separation only one break off point would exist and no seals would be necessary.

Figure 20.

AppendagesAppendages

Appendage items are:

- High gain antenna (HGA)

- Low gain antenna (LGA)

- Solar panels

- Attitude control jets

- Sun sensor and Canopus tracker

- High gain antenna (HGA)

The antenna dish will have a surface coating of white paint (a = 0.2, e = 0.9). The struts holding the antenna to the lander will be wrapped in a thermal blanket and the struts that are sunlit will be given a polished finish. The HGA has a wide operating temperature range and so the heaters will only be needed for the HGA actuators.

- Low gain antenna

A polished finish will be used for the center post and the reflector exterior. The interior of the reflector will be painted white.

- Solar panels

The backside of the panels will be painted white and the supporting structure of the sunlit side covered in a grahite/epoxy coating (a = 0.84, e = 0.85) These will only be deployed at Phobos.

- Attitude control jets

The attitude control jets on both the landers and the propulsion module will have a pattern of white paint and polish.

- Sun sensor and Canopus tracker

The sun sensor will be blanketed and shaded, and the sunlit side of the canopus tracker will be blanketed.



Thermal System Mass BreakdownThermal System Mass Breakdown

Dry mass of the spacecraft is 1134 kg. Thermal subsystem constitutes about 4% of this dry mass.
Mulitlayer Insulation 10 kg
4 SECs4 kg
Louvers8 kg
Pipe from RTG 16 kg
Coatings and paints 4 kg
Electrical heater 1 kg
Miscellaneous 2 kg
Total:45 kg

All relevant thermal control calculations are included in Appendix D.

Design Strengths and WeaknessesDesign Strengths and Weaknesses

One of the most significant strengths of this vehicle design is the elimination of the typical bus structure by the use of a modular design, the main advantage being less structural mass and more scientific payload capability. Other strengths include the use of two landers which allows a certain amount of redundancy in the event of failure of the mobile lander. The weak links in this design are non-redundant communication with Earth, vulnerability to micrometeoroid impact with the main tanks, and the possible risk of deployment problems with the sensor turrets.

References

Agrawal, Brij N., Design of Geosynchronous Spacecraft, Chapter 6: Electric Power, pp. 323-382, Prentice-Hall, Inc., Englewood Cliffs, New Jersey, 1986.

B. Kit and D.S. Evered, Rocket Propellant Handbook, MacMillian Co., New York, 1960.

George P. Sutton, Rocket Propulsion Elements: An Introduction to the Engineering of Rockets, Fifth Edition, John Wiley & Sons, New York, 1986.

Hering, R.G., "Thermophysics and Spacecraft Thermal Control," Progress in Astronautics and Aeronautics, Vol. 35, 1974.

Holmberg, N. A., Faust, R. P., Holt, H. M., "Viking '75 Spacecraft: Design and Test Summary, Vol I - Lander Design," NASA Reference Publication 1027, November 1980.

Holmberg, N. A., Faust, R. P., Holt, H. M., "Viking '75 Spacecraft: Design and Test Summary, Vol II - Orbiter Design," NASA Reference Publication 1027, November 1980.

Meyer, R., "Introduction to Space Technology Lecture Notes," UCLA Course MANE 161B, Winter Term 1991.

Moeckel, W.E., Propulsion Systems for Manned Exploration of the Solar System, Astronautics and aeronautics, Vol. 7, No. 8, 1969, pp. 66-77.

NASA News: Press Kit, Release No: 77-136, Project Voyagers 1 and 2, August, 1977.

P.G. Hill and C.R. Peterson, Mechanics and Thermodynamics of Propulsion, Addison-Wesley Publishing Company, pg. 491,1970.

Piellisch, Richard, New Solar Arrays Mean New Materials, Aerospace American, pp. 20-23, May 1991.

Space Vehicle Design, AIAA.

Spacecraft Mass Estimation, Relationships and Engine Data, Eagle Engineering Report #87-171, November 1987.

Surkov, Yu A. Exploration of Terrestrial Planets from Spacecraft: Instrumentation, Investigation, Interpretation. Ellis Horwood Limited, West Sussex, 1990.

Appendix A: Patched-Conic Analysis and Design Appendix A Patched-Conic Analysis and Design

The patched-conic analysis of a mission to Phobos is discussed here. To minimize the DV required, a Hohmann transfer orbit is used to take the spacecraft from the sphere of influence of the departure planet to the sphere of influence of the arrival planet. The energy of an Earth-Mars Hohmann transfer orbit is given by:


where m is the gravitational parameter of the sun, r1 is the Earth-Sun radius, and r2 is the Mars-Sun radius. The heliocentric departure speed, V1, required at the departure point is obtained from:


The difference between the heliocentric departure speed, V1, and the Earth's orbital speed, V, represents the speed of the spacecraft relative to the Earth upon exit from the Earth's sphere of influence. In other words, it is the hyperbolic excess speed left over after the spacecraft has escaped the Earth. Knowing that the orbital speed of the Earth is 1 AU/TU or 29.78 km/sec, V is found to be:


The correct phase angle at departure must exist for the spacecraft to encounter Mars at the time of arrival. The equation for the phase angle at departure is given by:


where n2-n1 is the difference in true anomaly at the two points, wt is the angular velocity of the Mars, and t2-t1 is the time-of-flight for the Hohmann transfer:




Therefore the correct phase angle at departure is 44.36 degrees which occurs on July 9, 2005. This date is chosen as the launch date for the mission. If this particular launch opportunity is missed, a waiting period of approximately two years is required for the same phase angle to repeat itself.

Once the hyperbolic excess speed is determined, it can be used to establish the injection or launch conditions near the surface of the Earth. Since energy is constant along the geocentric escape hyperbola, it is possible to equate e at injection and e at the edge of the sphere of influence where r = r.


With ro = 6678 km, Vo is found to be 11.32 km/sec. The circular speed of the spacecraft at a parking orbit altitude of 300 km is:


The difference between the injection speed, Vo, and the circular speed, Vcs, is the DV required to escape the sphere of influence of the Earth.


This speed change required to place the spacecraft on its 259-day transfer orbit is provided by a fourth-stage Titan IV rocket. During the transfer orbit, a series of mid-course corrections should be made to assure proper arrival at Mars. These corrections will amount to a total DV of 100 m/s. Since the orbit plane of Mars is inclined 1.85 degrees to the ecliptic, a mid-course plane change is also required during the transfer orbit. The geometry of this mid-course plane change is illustrated below.



















The plane change will occur at the node where the orbit plane of Mars crosses the ecliptic plane. The angular momentum, parameter, semi-major axis, and eccentricity of the transfer orbit can be computed from:





The radius and speed of the spacecraft at the time of the crossing can be determined from:



Therefore the DV required to produce a 1.85 degrees plane change is:


The next crucial maneuver after all course corrections have been accomplished will be that of entering orbit around Mars. It is assumed below that rocket braking alone will be used; later, aerobraking is described. The heliocentric arrival speed, V2 , and flight-path angle, f2 , can be determined from:



The hyperbolic excess velocity upon entrance to the sphere of influence of the Mars can now be obtained from the law of cosines:



where V is the orbital speed of Mars. The energy upon entrance to the sphere of influence with the energy at periapsis can be equated to get:


Assuming a periapsis radius of 9700 km, the periapsis velocity can be computed from the energy equation:



At the periapsis, the main engines will ignite to place the spacecraft into a circular orbit around Mars of 9700 km radius; at the same time, the orbit will be made equatorial. It should be noted that the plane of the Mars equator is inclined 23.98 degrees to the ecliptic plane. The geometry of the capture orbit and the plane change is shown below.
































The circular speed of the spacecraft at a radius of 9700 km is:


Therefore, the DV's required to circularize and equatorialize the orbit are:


The combined total DV can now be computed from:



This orbit is only 322 km away from the final destination of the spacecraft which is the orbit of the Phobos. The angular velocities of these circular orbits are:


Therefore, the synodic period between the spacecraft and the moon is found to be:


This means that a rendezvous with Phobos is possible almost every 6 days. The transfer to Phobos will be accomplished via a Hohmann transfer ellipse. The energy of the transfer orbit is :

We can now proceed to determine the speed change required to place the spacecraft on the transfer orbit.




The speed of the spacecraft upon arrival at the Phobos orbit is:


The DV required to insert the spacecraft into the lower circular orbit is:



The sum of DVto1 and DVto2 is the speed change required to transfer the spacecraft from the higher circular orbit to the lower circular orbit. This speed change will be called DV3.












Table A.1 sums up the total DV required for the mission.

Table A.1: Delta-V Summary (Rocket Braking Only)

_________________________________________________

DV

Classification (km/sec)

_________________________________________________

Mid-Course Trajectory Corrections 0.10

Plane Change at Crossing Node 0.82

Capture Orbit at Mars 2.70

Transfer Orbit to Phobos 0.04

Station Keeping and Landing 0.15

Miscellaneous 0.11

_________________________________________________

Total = 3.93

The total DV for the mission can be reduced if aerobraking is used to capture at Mars. The gains from using this technique are certainly attractive, but it is difficult to carry out. The approach sequence is illustrated below.
















The spacecraft has an initial hyperbolic excess speed of 2.65 km/sec. It aims for an elliptical orbit of 3405 x 9700 km. The guidance computer directs the craft for a closest approach of 25 km above the surface-well within the atmosphere. The speed at this point can be computed from the energy equation.


For the elliptical orbit (3405 x 9700 km), semi-major axis and speeds at periapsis and apoapsis are:




The difference between Vp and Vp is the speed loss that would result from plunging the spacecraft into the planet's atmosphere.


If all goes well, the craft zips out of the atmosphere into an orbit with the desired high point. On reaching the high point, the craft fires its engines for a speed boost, changing the shape of the orbit into a circular orbit. The speed of the circular orbit is:


Therefore the required speed change at the apoapsis would be:


As before, this speed change is combined with the speed change required to equatorialize the orbit to give a total DV of 1.40 km/sec. The remaining mission scenarios are the same as before. Table A.2 summarizes the mission DV's and points out the advantages of aerobraking.

Table A.2: Delta-V Summary (Aerobraking at Mars)

_________________________________________________

DV

Classification (km/sec)

_________________________________________________

Mid-Course Trajectory Corrections 0.10

Plane Change at Crossing Node 0.82

Capture Orbit at Mars 1.40

Aerocapture Attitude and Maneuvers 0.02

Orbit Trim at Apoapsis 0.03

Transfer Orbit to Phobos 0.04

Station Keeping and Landing 0.15

Miscellaneous 0.17

_________________________________________________

Total = 2.73

The penalty to be paid for the benefits of aerocapture is the increased complexity of techniques and systems. Also, further analysis is required to determine whether or not aerobraking at an altitude of 25 km would result in the desired speed loss of 1.37 km/sec. For these reasons, aerocapture is not considered for this mission.

Appendix B: Power DesignAppendix B Power Design

I. RTG Power design analysis

Power Requirement = 200 W

# of modules = 200 W .

20.5 W/module

" = 10 modules

Net weight = 10 x 2.2 kg

" " = 22 kg

Dimensions per = 2" x 18"

module

Net dimensions = 20" x 18"

Heat produced = 1800 W

by modules

II. Solar Array Design (GaAs cells)

Array voltage = 27.5 V x 1.2

= 33 V

Total power (EOL) = 200 W + 30.9A-h x 33V

15 h

= 268 W

Temperature effects = (50-28) x .0025

= .055

Array capacity = 268 W .

(1-.15)(cos5o)(1-.055)

= 335 W

Total cell area = 335 W .

600 W/m2 x 0.18

= 3.1 m2

# cells = 3.1/8x10-4

= 3875 cells

Array size = 3.1/0.9

= 3.4 m2

III. Battery Design (NiH2)

Capacity = 200 W x 4 hrs

28.75 x 0.9

= 30.9 A-h

ED = 20 W-h/lb

MB = 30.9 x 28.75

20

= 44.4 lbs.

= 20.0kgAppendix C : Main Propellant Tank Sizing

Propellant Data

Fuel: Monomethyl Hydrazine (MMH)

Oxidizer: Nitrogen Tetroxide (N2O4)

Propellants were chosen because of space storability and

fairly high Isp.

rN2O4 = 1447 kg/m3

rMMH = 878.8 kg/m3

Operating Temperature Range = 2°C - 100°C

Isp = 280 sec

C* = Exhaust Velocity = 1747 m/sec

Propellant Mass

O/F ratio = 2.17 (68.5% N2O4 , 31.5% MMH by mass)

Total main propellant for mission = 2125.5 kg

N2O4 mass = 0.685 (2125.5) = 1456.0 kg

MMH mass = 0.315 (2125.5) = 669.5 kg

N2O4 volume = 1456 kg / 1447 kg/m3 = 1.006 m3

MMH volume = 669.5 kg / 878.8 kg/m3 = 0.762 m3

Spherical Tank Sizing

Assume 5% ullage volume : MMH Vol. = 1.056 m3

N2O4 Vol. = 0.800 m3

Volume of Sphere = 1/6*p*diameter3

For 2 tanks (1 MMH,1 N2O4): dMMH = 1.263 m

dN2O4 = 1.152 m

For 4 tanks (2 MMH,1 N2O4): dMMH = 1.003 m

dN2O4 = 0.914 m

Tank Wall Thickness (4 tank option)

-Titanium tanks (Ti-3Al-13V-11Cr)

-1000 psi nominal tank pressure (w/ safety factor of 1.2

tank pressure = 1200psi)

- stit = yield strength of titanium = 135 ksi = 931 Mpa

tMMH = p(d/2)/2(s) = (1200)(0.5015)/2(135) = 2.23 mm

tN2O4 = 2.03 mm

Tank Dry Mass

rtitanium = 4400 kg/m3

Volume of titanium for MMH tanks =

(2 tanks) x [p/6(d1 +2t)3] - [p/6(d1)3] = 0.01415 m3

Volume of titanium for N2O4 tanks = 0.010702 m3

Mass of MMH tanks = rV = (4400)(0.01415) = 62.3 kg

Mass of N2O4 tanks = rV = (4400)(0.010702) = 47.1 kg

Total main propellant tank mass = 109.4 kg

Appendix D: Thermal Analysis and CalculationsAppendix D Thermal Analysis and Calculations

Thermal balance equation : mc(dT/dt) = asIsAs + P - eAsT4

as = Average solar absorptance of spacecraft Is = Solar radiation flux per unit area

As = Projected area as seen by sun P = Equipment heat

e = Average emissivity A = Total surface area of spacecraft

s = Stefan-Boltzmann constant T = Spacecraft temperature

Calculating Is : Is = 3.86 x 1023/ 4pd2 d = distance of spacecraft from sun

Is at Earth = 1362 W/m2 Is at Mars = 591 W/m2

Graph below shows variation of solar flux radiation versus distance























Calculating A and As :

For calculating the areas a very approximate method was used.

The height of the vehicle is 5.5m and the width is 3m. Assuming the spacecraft is a triangle, the area is 1/2bh.

Area of side of spacecraft = 7.5m2

Assuming the spacecraft flies with the propulsion module facing the sun, the area as seen by the sun is 3m x 3m = 9m2

Area as seen by the sun is 9m2 + extra(10%)= 10m2

As = 10m2

Total surface area is area of top plus four sides plus bottom.

A= (1.2)2 + 4(7.5) + 10 = 41m2

A = 41 m2

Calculation of equipment heat, P:

heat from instruments = 20 W

RTG power is 200 W, efficiency is 90% so heat from RTG = 1800 W

Total heat, P = 1800 + 20 = 1820 W

P = 1820 W

Now from thermal balance at steady-state dT/dt = 0 so

TE = (( asIsAs + P)/eAs)1/4

Calculating as and e :

as = 1/As S asj Asj

e = 1/A S e k Ak

as : Aluminized Kapton blanketing, a = 0.35 75% of 10 m2

Black paint on boosters, a = 0.93 25% of 10 m2

As = 10m2

therefore as = 1/10((7.5 x 0.35) + (2.5 x 0.93)) = 0.495

as = 0.495

e : Aml (mobile lander) = 4.5 m2

Asl (stationary lander) = 5.5 m2

Apm (propulsion module) = 30 m2

Aml : 15% louvers, 85% rest

Asl : 15% louvers, 85% rest

Apm : 15% boosters, 25% blanket, 60% aluminum foil

louvers, white paint e = 0.88

landers are painted with aluminum silicone paint e = 0.28

boosters, black paint e = 0.88

MLI aluminum foil e = 0.04

MLI e = 0.6

e = 1/40((0.15 x 4.5 x 0.88) + (0.85 x 4.5 x 0.28) + (0.15 x 5.5 x 0.88)

+ (0.85 x 5.5 x 0.28) + ( 0.15 x 30 x 0.88) + (0.25 x 30 x 0.6)

+ (0.6 x 30 x 0.04)) = 0.322

e = 0.322

Now using TE = (( asIsAs + P)/eAs)1/4 we get:

TE at Mars = 283.9°K = 10.9°C

TE at Earth = 329°K = 56°C

These figures are a very rough approximation, however they do provide a reasonable range for the spacecraft to operate within.

Appendix E: Spacecraft Drawings





Phobos Mission Vehicle Design







Submitted to Dr. Wallace Fowler in Partial Fulfillment of the Requirements of ASE 396 Space Systems Design.









by

Guillaume Brat - Communication

Mustafa Khan - Thermal

Jose Lozano - Power

Walter L. Moguel - Structures

Kamran Moosavi - Trajectory

David North - Propulsion

Harold E. Pangilinan - Scientific Sensors








December 10, 1991

Abstract

The preliminary design of an unmanned probe vehicle to the Mars moon Phobos has been developed. The primary goal of this mission is to gather chemical composition data with a particular emphasis on the search for water. Secondary goals are to completely map the surface of Phobos and investigate its internal structure via seismographic experiments. The mission will be accomplished with a three-module spacecraft including a propulsion module, a stationary lander, and a smaller mobile lander. Launch from Earth will be made with a Titan IV/Centaur G with a nominal launch date of July 9, 2009 and a total mission time of approximately one year. Preliminary design of vehicle subsystems includes propulsion, power, communication, scientific sensors, structures, thermal control, guidance, navigation and control. In addition an analysis of trajectories has been made.