Spacecraft Design Archive










Talia Jurgens




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Current System Analysis
Planetary Mission Analysis Using Microspacecraft Technology Systems

Tony Statom

November 2, 1992

The Jet Propulsion Laboratory using the developments in microtechnology from DARPA and SDIO are investigating the potential use of this technology in Microspacecraft.

The history of spacecraft shows that from the small mass spacecraft that the space program began with the mass has steadily increased since then. The increased mass of spacecraft are causing mission frequency to decrease. This is the driver for the development of spacecraft using microtechnology. The decreased mass should increase launch frequency.

Definitions for spacecraft are as follows[3]:
Standard1000 kg
Small100 kg
Microspacecraft10 kg

Microspacecraft Workshop Summarization
In July of 1988 NASA and SDIO sponsored a workshop at JPL titled "Microspacecraft for Space Science". The results are[2]:

  1. Microspacecraft (1-10 kg) are technically feasible.
  2. There is a class of scientific and exploration missions that can be enabled by microspacecraft. This class of missions requires many simultaneous measurements displaced in position, as on the surface of a planet or small body or in the region of space. The enabling feature of microspacecraft is the assertion that using many microspacecraft (1-10 kg) will cost less (spacecraft and launch costs) and involve less risk than using large (500-1000 kg) spacecraft for such missions.

  3. Other missions enabled by the microspacecraft concept are those that require very high mission delta-V's.
  4. While useful and perhaps enabling for the types of missions mentioned above, microspacecraft are not applicable to all types of space exploration and science missions and should not be viewed as a panacea.

Microspacecraft Technology
The SDIO and DARPA programs have yielded developments that are being used in conceptual designs including most spacecraft subsystems.

Solar arrays or nuclear (radioisotope) sources can be used to generate power. A 1 x 2 meter inflated solar array that can be stored in a 10.2 cm diameter and 1 m long canister[2]. The beginning life power is 125 watts and 3 year end of life power of 100 watts with a mass of 0.66 kg[2]. A microradioisotope thermal electric generator (RTG) with a mass of 159 grams produces 2 watts at 5 volts after 5 years[2].

One of the objectives from the SDIO investigation was small projectiles with excellent maneuverability. A small propulsion subsystem with a mass of around 50 grams each using standard bi-propellants can be used for main engines and 4 small engines for attitude control thrusters with a mass of less than 3 kg an rated over 600 N each[2]. SDIO also contributed to the development of a small upper stage called the Advanced Liquid Axial stage (ALAS). Specific impulse about 345 seconds. Wet mass of around 7.7 kg and dry of about 1.8 kg[2].

A small unit developed by the SDIO program was a system with mass of 150 grams called Quartz Rate Sensor (QRS). The inertial rates are measured by the quartz tuning fork elements, and linear accelerations by the silicon accelerometers. This system was made by Systron Donner[2]. It draws 7 watts of peak power.

The Lawerence Livermore National Laboratory has developed a miniature star tracker camera the prototypes mass is about 300 grams[2].

The developments in Scanning Tunnelling Microscopy (STM) technology allows the use of the quantum-mechanical electron tunnelling to be used for position detection. This sensor would be small in size. The sensor is a few Angstroms[5]. This sensor can also be used for a Martian microseismometer and microweather station[2].

A parallel processor using two dimensional hybrid wafer scale integration has a mass of about 3.6 kg is being developed by The Space Computer Corp. This processor will have 1.3 giga floating point operations per second. A future version using three dimensional hybrid wafer scale integration is expected to have a mass of about 0.8 kg[2].

A conceptual design using optical frequencies has been done. It uses a semiconductor laser at a wavelength of 8.0 x 10-7 m. The subsystem uses about 6 watts of power, a 10 cm aperture, and should weigh about 1 kg. Performance at night is 1000 bits per second from 1 AU to a 10 m diameter receiver located Earth with clear skies. With daytime background noise the performance drops to 1 bit per second[2].

A conceptual design for a camera has a mass of less than a kilogram uses 4 watts in operation and has a resolution of 7 meters per pixel at a range of 100 km. The design assumed a spinning rate of 5 rpm or less. The system included a data buffering which caused a read out rate to the microspacecraft of about 10 bits per second[2,6].

Using microspacecraft gives greater options in the choice of a launch vehicle. With these small systems greater delta V can be obtained. A microspacecraft launched at 50 km/s would reach 10 ,100 ,1000 AU from the Sun in 0.6, 6.2 and 62.8 years respectively. This velocity might be obtained in the future by electromagnetic launchers[4]. If conventional chemical launch vehicles are used and a microspacecraft is launched at 10 km/s then 10, 100 and 1000 AU from the Sun can be reached in 2.2, 34 and 390 years respectively[4]. This assumes:

  1. Impulsive delta V of between 10 and 50 km/s.
  2. The payload is launched from a 500 km circular Earth orbit.
  3. The payload is launched at the right time and in the right direction in order to take advantage of its orbital energy relative to both the Earth and the Sun.
  4. The payload is launched on a heliocentric, hyperbolic escape trajectory[4].

Missions/Systems Descriptions
At JPL the three areas of focus in the current microspacecraft effort are:

  1. Indentifying technology needs for microspacecraft and micro instruments.
  2. Bringing the Asteroid Investigation with Microspacecraft (AIM) concept to the next level of feasibility by backing off on some technology assumptions, developing more detail it the spacecraft design, building a full-scale mockup of the spacecraft, and holding technical peer reviews.
  3. New conceptual studies to increase the number of possible missions considered for microspacecraft[1].

Astroid Investigation With Microspacecraft (AIM)
The mission objective is to flyby three separate near Earth asteroids and return high resolution images[1]. Constraints:

  1. Sun range between 0.8-1.2 AU
  2. Earth-S/C range<1.6 AU
  3. Post launch delta V <200m/s [1].

Some subsystem descriptions:

  1. Launch vehicle - Pegasus carrying three microspacecraft[1].
  2. Telecommunications - can achieve 100 bps at 1.6 AU to the 34 m HEF ground stations of the DSN[1].
  3. Power - GaAs/Ge solar cells with a LiTiS2 secondary battery supplement[1].
  4. Attitude Control - three axis using Nitrogen pressurant form the hydrazine propulsion system and 0.005 lbf thrusters perform precise maneuvers[1].
  5. Command and Data - VLSI 1750A processor with 3 MIPS[1].
  6. Structure - GrEp honeycomb with aluminum face sheets for thermal conductivity. The shape is six sided or a hexagonal cylinder.

Mars Rover Sample Return (MRSR)
The mission objective is basically to return a sample from mars.

Subsystem descriptions[1]:

  1. Launch Vehicle - Atlas IIAS.
  2. Spacecraft Description
    1. the aerocraft
    2. the lander (Sample Return Capsule (SRC))
    3. the Mars Ascent and Return Vehicle (MARV)
    4. the rover
    5. the Sample Return Capsule (SRC)
  3. Thermal - Quartz nitro phenolic material for heat shield.
  4. Structure - Aluminum lithium.
  5. Power - LiSO2 battery for 0.5 W RF beacon transmitter. Other subsystems use LiTiS2 batteries and Silicon solar cells.
  6. Attitude Control - small fiber optic gyro and a JPL designed CCD line array star scanner.

Venus Atmospheric Sounder
The objective of the mission is to obtain full coverage of the atmosphere.

Some subsystem description[1]:

  1. Attitude Control - Small reaction wheels in pitch and roll only achieve three axis pointing control.
  2. Data System - Utilizes a VLSI 1750A processor and provides 1.1 Gbit of storage.
  3. Power - GaAs solar cells and LiTiS2 battery during occultations.
  4. Launch System - Taurus.

Comet Nucleus Mission
The objective of this mission is to return a sample from a comet.

Some of the subsystems[1]:

  1. Power - LiTiS2 Battery and a half size RTG (as opposed to a single GPHS RTG used on Galileo)
  2. Propulsion - NTO/N2H4 bipropellants stored in carbon-wound titanium tanks. Utilizes a Rhenium chamber for the main engine that has an Isp of 328 sec.
  3. Command and Data - 1750A processor. Downlinks through a Ka-band link using a planar phased array antenna.

Summary and Conclusion
With the advances in microtechnology applicable to spacecraft through the developments of DARPA and SDIO microscpacecraft conceptual designs are being analyzed. The microspacecraft gives greater flexibility in the choice of a launch vehicle. This flexibility should drive down mission costs and increase launch frequency. While there are many missions that can utilize microspacecraft it is not viewed as all inclusive.

The microspacecraft analysis is not complete. Greater fidelity to the analysis both technically and economically must be done. Subsystem component options should be narrowed. Power sources with long duration must be developed to return data over interplanetary distances.

This type of approach to spacecraft system will continue as microtechnology makes greater strides in research. As microtechnology continues to decrease in size and expand in ability it should be considered in initial conceptual design to assess feasibility.

Ross Jones the supervisor of the Advanced Spacecraft System Concepts Group at JPL who sent me three of his papers that this system analysis was based on.


  1. Short, L.P. et. al. "Planetary Missions Using Microspacecraft Technology", IAF-92-0821, 43rd Congress of the International Astronautical Federation, Washington,DC, August 28-September 5, 1992.
  2. Jones, R.M., and Salvo, C.G. "Microspacecraft Technology For Planetary Science Missions", IAF-91-051, 42nd Congress of the International Astronautical Federation, Montreal, Canada, October 5-11, 1991.
  3. Jones, R.M., "Small Spacecraft Activities at JPL", Utah State University Conference on Small Satellites, August 26, 1991.
  4. Jones, R.M., "Microspacecraft Missions and Systems", Journal of the British Interplanetary Society, Vol. 42, #10, p 448, October 1989.
  5. Waltman, S.B., and Kaiser, W.J., "Electron Tunnel Sensor Technology", Journal of the British Interplanetary Society, Vol. 42, p 474, October 1989.
  6. Ravine, M.A., and Soulanille, T.A., "Cameras For Microspacecraft", Journal of the British Interplanetary Society, Vol. $2, p 460, October 1989.