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THE LIFETIME OF THE EXOSAT SPACECRAFT

1. Background

EXOSAT's design specification was based on a two-year operational phase. This had a number of major influences on the system design, particularly on the requirements for orbit stability, end-of-life solar array power and attitude/manoeuvre fuel.

This article examines the actual situation after 21 months of operation and considers those factors which will determine the ultimate lifetime expectancy of the mission.

2. Orbit Considerations

The EXOSAT orbit had to satisfy the following criteria:

  • maximum time spent outside the Van Allen radiation belts.
  • line of apsides (apogee-perigee line) close to perpendicular to the ecliptic plane, to maximise the area of sky over which lunar occultations could be effected.
  • a Northern apogee such that a single, Northern hemisphere ground station (Villafranca, Spain) could be used for telemetry reception, telecommanding and tracking.
These criteria led to a choice of orbit with the nominal characteristics as shown in table 1.

The constraints on the time of day of launch for 26th May 1983 are shown in Fig. 1 and were:

  • the Solar Aspect Angle (SAA) during the initial transfer orbit had to lie between 30* and 120%
  • there should be no eclipse during the transfer orbit.
  • the final orbit lifetime had to be greater than 2 years.
  • the maximum eclipse duration in the final orbit had to be less than 60 minutes from considerations of battery design.
The EXOSAT launch took place 1 minute prior to window closure and the orbit achieved had the characteristics as shown in Table 1 and a lifetime of approximately 3 years.

TABLE 1

Orbital Elements Nominal Achieved
(26.5.83)

Height of Perigee 350 km346.6 km
Height of Apogee 196235 km191708.7 km
Eccentricity 0.93550.93433
Inclination 72.5072.470
Argument of Perigee 286.50286.360
Orbit Period 93.6 hrs 90.6 hrs

EXOSAT's orbit is perturbed strongly by the gravitational influence of the Sun and the Moon and this has the (principal) effect of changing the eccentricity of the orbit whilst the semimajor axis remains essentially constant. Thus the perigee height has increased from 350 km initially to a maximum of ca. 4500 km in November 1984, since when it has gradually decreased such that without active control, the EXOSAT spacecraft will re-enter the Earth's atmosphere during the second-half of April 1986 (see Fig. 2).

EXOSAT has, however, a hydrazine reaction control system with a capability of imparting 173 m/sec to the orbit. This system was included to permit active orbit control in order to modify the orbit period by.small increments to achieve the geometry required for moon occultation of Xray sources. Hydrazine usage to date has been 2.3 m/sec for commissioning and calibrating the system.

If the remaining 170 m/sec were used for perigee-raising, the lifetime of the orbit could be extended by about 1 year ie. until April 1987. The amount of the extension increases slightly the later the orbit manoeuvres are delayed; this effect is demonstrated in Figures 3 to 6. The optimal strategy is to perform small manoeuvres starting in April 1986 such that the perigee height is maintained just above the critical value at which the orbit would decay due to atmospheric friction (a few hundred kilometres). Other considerations, however, will probably dictate that less than the full 170 m/sec is used to extend the lifetime (see Section 5 below).

3. Power from the Solar Array

With the exception of eclipses (when battery power is used) the subsystems and experiments are powered from the Solar Array (consisting of 3312 solar cells) mounted on top of the spacecraft and rotated such that it remains near-normal to the Sun direction.

The output from the Array varies because of:

  1. degradation of the solar cells from aging and irradiation.
  2. variation in the Solar input caused by the eccentricity of the Earth's orbit around the Sun. This effect produces a variation of ca. 7% in Solar input from perihelion (2nd January) to aphelion.
The superposition of these two effects since launch can be seen in Fig. 7, which shows the Array output as derived from the Sun bus current and voltage sensors. The "noise" on this curve is the combined effect of telemetry quantisation, Array offsets from normal (up to 5°) and occasional high readings taken close to perigee when there is a significant input to the Array from Earth albedo. The power currently required from the Array is around 230 Watts. This is not a constant load figure but includes also the "peaks" which occur as a result of aperiodic operation of thrusters and the temperature control circuits of the gyros and therefore represents the level at which battery discharge will not occur. From the trend of the Array output to date, it can be seen that this load can be supported throughout the remaining natural mission lifetime and any possible extension.

4. Attitude Control Fuel

Limit-cycling and attitude manoeuvres (slews) are achieved using a cold-gas propane reaction control system. 14 kilograms of propane were loaded, with a prelaunch allocation between attitude manoeuvres and attitude stabilisation of approximately 55%:45%. This manoeuvre allocation was based on a (nominal) slew rate of 300 deg/hr. In practice, the amount of fuel used for manoeuvres has been somewhat less, since the slew rates used have been lower (171 deg/hr was eliminated in Dec. 1984). However a number of anomalies which have occurred with the Attitude and Orbit Control System, (AOCS, see earlier reports) have used additional propane fuel, for which there was no allowance in the original budget.

Although there is no direct measure of either the amount of propane used or the amount remaining in the tanks, two indirect but somewhat inaccurate methods are available:

  1. Propane logging. Measurements have been made of the fuel consumption for limit-cycling and manoeuvring. This involves an operational procedure whereby the commanded propane plenum pressure is reduced for a short time and the rate of fall of the pressure is monitored. A housekeeping exercise is then carried out to integrate the consumption over time, adding in separately estimates for the additional fuel used as a result of anomalies (spurious triggering of Safety Mode, thrusters "stuck-on" etc). This method indicates that approximately 5.5 kg of propane remains as of-the middle of February 1985.
  2. Propane gauging.The propane tanks are equipped with heaters with a total rating of 2.4 Watts, which are normally switched on only during eclipses. Switching on these-heaters and monitoring the subsequent temperature rise (equilibrium is reached after about 4 days) allows an estimate of the propane remaining in the tanks.
    Such an exercise was undertaken in the middle of February 1985 and the preliminary results give a "best fit" to a remaining propane mass of 4 kg.
These exercises are currently being re-examined in an attempt to reconcile the substantially differing results. Meanwhile, two measures are being applied to reduce the rate of consumption:

  1. reduction of the manoeuvre frequency during the forthcoming A03 programme, ie. increase the mean observation duration, and limitation of the slew rate to 42.7 deg/hr.
  2. Use of the Fine Sun Sensor for roll control (with the OBC program SMC compensating for sun motion) whenever possible. This measure results in lower noise injection into the X-axis limit cycle than the alternative 2-star roll control technique.
In this manner the overall consumption can be reduced to just over 2 kg per year for future operations, assuming of course that no further fuel-consuming anomalies occur.

Analysis of the calibration orbit manoeuvre indicates that, if all of the 170 m/sec were utilised, this could cost up to 1 k9 of propane for attitude stabilisation during the orbit manoeuvre. This fuel is required to compensate for the misalignment between the line of action of the hydrazine thruster and the spacecraft centre of mass.

5. Conclusions

The optimum use of the remaining propane will have been made if the spacecraft re-enters the atmosphere at exactly the same time as the propane is exhausted.

One method of achieving this would be to leave the orbit uncontrolled until April 1986 and then to apply small orbit corrections to extend the life until such time as the propane is exhausted.

The approach however has two risk factors associated with it:

  1. with only a single actuation to date, there is not yet a high confidence associated with the operation of the hydrazine system. Failure of a single manoeuvre late in the mission could be irrecoverable.
  2. it is not possible to perform orbit manoeuvres unless there are 3 operable gyros available to control the attitude. The X-gyro spin motor current rose dramatically on day 366 1984 and has displayed an erratic behaviour since. In the middle of February, it rose again and caused a significant temperature increase of the gyro box as a result of which the X-gyro was switched off. Prior to this step however, there was good reason to believe that the X-gyro could still be used to control the spacecraft. If a second gyro fails, it may be possible to "resurrect" the X-gyro to reestablish 3-axis gyro control.
All factors considered, the current philosophy is to leave the orbit uncontrolled until at least early 1986. If the anomalous consumption of propane has then been minimal throughout 1985, the strategy can be reviewed in the light of the best estimate of the propane remaining at that time.

A. Parkes


Fig.1 EXOSAT Launch Window
exosat launch window
Fig.2 Nominal orbit without control
Fig 2. perigee versus exosat lifetime (description above)
Fig.3 Controlled orbit. Sub-optimal strategy. Full perigee raising is done in March 1985.
Fig 3. perigee versus exosat lifetime (description above)
Fig.4 Controlled orbit. Sub-optimal strategy. Full perigee raising is done in September 1985.
Fig 4. perigee versus exosat lifetime (description above)
Fig.5 Controlled orbit. Sub-optimal strategy. Full perigee raising is done in April 1986.
Fig 5. perigee versus exosat lifetime (description above)
Fig.6 Superposition of Figs 2, 3, 4, and 5.
Fig 6. perigee versus exosat lifetime (description above)
Fig 7 superposition of two effects (description in main text)
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