COBE Mission Design, Spacecraft and Orbit

This summary is an excerpt from Bennett (1992, Proceedings of the Third Teton Summer School: The Evolution of Galaxies and Their Environment, H. A. Thronson and J. M. Shull eds.); more detailed information is given in references cited therein. For more information on this subject please see Boggess et al. (1992).

The need to control and measure potential systematic errors led to the requirements for an all-sky survey and a minimum time in orbit of six months. The instruments required temperature stability to maintain gain and offset stability, and a high level of cleanliness to reduce the entry of stray light and thermal emission from particulates. The control of systematic errors in the measurement of the cosmic microwave background anisotropy and the need for measuring the interplanetary dust cloud at different solar elongation angles for subsequent modeling required that the satellite rotate.

In near-Earth orbit, the Sun and Earth are the primary continuous sources of thermal emission and it was necessary to ensure that neither the instruments nor the dewar were exposed to their radiation. A circular Sun-syncronous orbit satisfied these requirements. An inclination of 99 deg and an altitude of 900 km were chosen so that the orbital plane precesses 360 deg in one year due to the Earth's gravitational quadrupole moment. The 900 km altitude is a good compromise between contamination from the Earth's residual atmosphere, which increases at lower altitude, and interference due to charged particles in the Earth's radiation belts at higher altitudes. A 6 PM ascending node was chosen for the COBE orbital plane; this node follows the terminator (the boundary between sunlight and darkness on the Earth) throughout the year. By maintaining the spacecraft spin axis at about 94 deg from the Sun and close to the local zenith, it is possible to keep the Sun and Earth below the plane of the instrument apertures for most of the year. However, since the Earth's axis is tilted 23.5 deg from the ecliptic pole, the angle between the plane of the COBE's orbit and the ecliptic plane varies through the seasons from -14.5 deg to +32.5 deg. As a consequence, the combination of the tilt of the Earth's axis, the orbit inclination, and the offset of the spacecraft spin axis from the Sun brings the Earth limb above the instrument aperture plane for up to 20 minutes per orbit near the June solstice. During this period the limb of the Earth rises a few degrees above the aperture plane for part of each orbit, while on the opposite side of the orbit the spacecraft goes into the Earth's shadow. In the nominal COBE orbit, the spacecraft's central axis scans the full sky, though not with uniform coverage, every six months. The orbital period is 103 minutes, giving 14 orbits per day.

A 3-axis attitude control system was implemented by using a pair of inertia wheels (yaw angular momentum wheels), with their axes oriented along the spacecraft spin axis. These wheels carry an angular momentum opposite that due to the spacecraft rotation to create a nearly zero net angular momentum system. The spacecraft orientation is controlled by three reaction wheels with spin axes 120 deg apart in the plane perpendicular to the spacecraft spin axis and by electromagnetic coils (torquer bars) that interact with the Earth's magnetic field. Earth and Sun sensors (one of each on each of the
three transverse control axes) provide control signals to point the spin axis away from the Earth and at least 90 deg from the Sun. Rate damping and fine resolution attitude sensing are provided by six gyros, one on each transverse control axis and three on the spin axis. Coarse attitude parameters are calculated by using telemetered data from the attitude control sensors to produce attitude solutions good to 4 arcmin (1 sigma). A fine aspect is determined by using gyro data to interpolate between the positions of known stars detected in the short wavelength bands of the DIRBE instrument. The fine aspect solution has an accuracy of 1.5 arcmin (1 sigma) and is used in the analysis of data from all three instruments.

The FIRAS instrument, located inside the dewar, points along the spin axis with its 7 deg field of view. The three pairs of DMR receivers are spaced 120 deg apart around the aperture plane of the dewar. Each radiometer channel measures the difference in sky signal from a pair of horn antennas defining 7 deg fields of view separated by 60 deg, each beam being 30 deg from the spin axis. The spin causes a short-term interchange of the two beams associated with a single differential radiometer and thereby gives a modulation of the differential sky signal at the spin rate. The 0.8 rpm spin rate was chosen to be fast enough to reduce the noise and systematic errors that could otherwise arise from radiometer gain and offset instabilities. The DIRBE, also located inside the dewar, views 30 deg from the spin axis. The spin allows DIRBE to measure the emission and scattering by the interplanetary dust cloud over a range of solar elongation angles for each celestial direction, which aids in the discrimination and subsequent modeling of zodiacal radiation. DMR and DIRBE trace out a pattern of epicycles that enable them to scan half of the sky every day and obtain multiple measurements for each pixel of the sky.

The dewar is a 650 liter superfluid helium cryostat that kept the FIRAS and DIRBE instruments cooled to about 1.6 K. A deployable dewar aperture cover protected the cryogen and permitted calibration and performance testing of the cryogenic instruments prior to launch. A contamination shield attached to the inside of the dewar cover protected the DIRBE primary mirror from particulate or gaseous contamination until ejection of the dewar cover in orbit. It also protected DIRBE from emission from warm parts of the cryostat during ground testing. The depolyed conical Sun-Earth shield protects the scientific instruments from direct solar and terrestrial radiation and provides thermal isolation for the dewar. The shield also provides the instruments isolation from Earth-based radio frequency interference (RFI) and from the spacecraft transmitting antenna. The shield was designed to be flexible and was folded to fit within the Delta rocket fairing for launch. Contamination covers attached to the Sun-Earth shield were placed over the DMR horn antennas and were pulled away in orbit by the deployment of the shield. The deployed solar arrays provide the nominal spacecraft and instrument power load of 542 Watts.

The COBE has two omnidirectional antennas, one to communicate with the Tracking and Data Relay Satellite System (TDRSS), and the other to transmit data stored on tape recorders directly to the ground. The antennas are located on a mast at the bottom of the spacecraft deployed after launch. The COBE has a command and data handling system that stores and decodes the commands received from the ground, collects data from the instruments and spacecraft at the rate of 4 kbps, and prepares data for transmission to the ground. The on-board tape recorders and data system allow 24 hours of data to be transmitted to the Wallops Flight Facility in a single 9 minute pass. The data rate allocations for DIRBE, FIRAS, and DMR are 1716, 1362, and 250 bps, respectively. The remainder of the telemetry is assigned to spacecraft subsystems.

The COBE, as initially proposed, was to have been launched by a Delta rocket. However, once the design was underway, the Shuttle was adopted as the NASA standard launch vehicle. After the Challenger accident occurred in 1986, ending plans for Shuttle launches from the West Coast, the spacecraft was redesigned to fit within the weight and size constraints of the Delta. The final COBE satellite had a total mass of 2,270 kg, a length of 5.49 m, and a diameter of 2.44 m with Sun-Earth shield and solar panels folded (8.53 m with the solar panels deployed).

Ground testing of the COBE was necessary to demonstrate that the individual sub-systems, and ultimately the entire spacecraft and instrument assembly in its flight configuration, could satisfy both the sensitivity and systematic error requirements. System level tests were performed: to simulate the space environments, including vacuum and temperature; to determine susceptibility to vibration, acoustic excitation, and acoustic shock; to quantify electromagnetic interference (EMI) self-compatibility and RFI susceptibility; to determine the interaction between instruments and spacecraft; to simulate the thermal and power conditions that would occur during eclipse periods; to test the deployables and moving parts (Sun-Earth shield, antenna boom, solar panels, dewar cover, FIRAS external calibrator and moving mirror transport, and DIRBE shutter and chopper); and characterize and calibrate the instruments.

The COBE was launched aboard Delta rocket No. 189 at 1434 UT on November 18, 1989 from the Western Space and Missile Center at Vandenberg Air Force Base, California. The DMR receivers began operating the day after launch. The dewar cover was ejected three days after launch, and the FIRAS and DIRBE instruments began obtaining data on the same day. During the first month in orbit, various tests were undertaken to evaluate the performance of the instruments and spacecraft, and to optimize instrument parameters.

The COBE operated in a routine survey mode. The three instruments completed their first full sky coverage by mid-June 1990, and returned high quality data until the depletion of the liquid helium at 0936 UT September 21, 1990. The FIRAS, which had surveyed the sky 1.6 times, ceased operating when the helium ran out, but the DMR is still operating normally in all of its six channels. By November 1991 (over one year after helium depletion) the dewar temperature at the DIRBE detectors was about 50 K. The six longest wavelength bands were turned off in September 1990, but the four short wavelength bands of the DIRBE continue to acquire data at reduced sensitivity. The detector system responsivity in the short wavelength bands decreased by about an order of magnitude following cryogen depletion (largely due to the change in load resistance). However, sky maps of the large scale interplanetary dust signals are of adequate quality to permit searching for evidence of temporal changes on annual time scales.

In flight, the helium temperature inside the main cryogen tank was 1.40 K and the temperature of the inner surface of the Sun-Earth shield was 180 K. As expected, the Earth limb rose a few degrees above the Sun-Earth shield for a part of every orbit during a three month period starting in May. At these times, the Earth's radiation produced thermal transients in the instruments and adversely affected data for a portion of each orbit. Some of these data are still usable after careful calibration. One of the gyros for a transverse control axis failed electrically on the fourth day after launch. On September 7, 1991, one of the three gyros on the spin axis failed, but no data were lost and satellite operations continue in the nominal orbit.

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