This leads to the conclusion that anisotropically rejected thermal radiation cannot be ignored when we evaluate the evolution of the Pioneer 10 and 11 trajectories, and must be accounted for with as much precision as possible.
Far from the Sun and planets, the only notable heat sources on board Pioneer 10 and 11 are internal to the spacecraft. We enumerate the following heat sources:
While strictly speaking it is not a heat source, one must also consider microwave radiation from the spacecraft’s radio transmitter and HGA, as this radiation also removes what would otherwise appear as thermal energy from within the spacecraft.
The RTGs are the most substantial sources of heat on board. Each of the four generators on board the spacecraft produced 650 W of power at the time of launch, of which 40 W was converted into electrical energy; the rest was radiated into space as waste heat. The radiation pattern of the RTGs is determined by the shape and composition of their radiating fins (see Figure 2.5), but, notwithstanding the possible effects of aging, it is believed to be fore-aft symmetrical.
|RTG#||Spacecraft||Location||Test date||Thermal power (W)|
|44||Pioneer 10||Outboard||Oct 1971||649|
|45||Pioneer 10||Inboard||Nov 1971||646|
|46||Pioneer 10||Outboard||Nov 1971||647|
|48||Pioneer 10||Inboard||Dec 1971||649|
|49||Pioneer 11||Outboard||Sep 1972||649|
|51||Pioneer 11||Inboard||Oct 1972||650|
|52||Pioneer 11||Outboard||Oct 1972||649|
|53||Pioneer 11||Inboard||Oct 1972||649|
The amount of power generated by the RTGs is determined by the radioactive decay of the 238Pu fuel on board. The radioactive half-life of the fuel is precisely known ( 87.74 years), and the total power output of each RTG was measured before launch (Table 2.4).
The amount of power removed from each RTG in the form of electrical energy is telemetered to the ground. (Specifically, the RTG output voltage and current for each individual RTG is telemetered.) The remainder of the RTG power is radiated in the form of waste heat.
For each RTG, two temperature measurements are also available. One sensor, internal to the RTG, measures the temperature at the hot end of the bimetallic thermocouples. The other sensor measures the temperature near the root of one of the RTG radiating fins.
Next to the RTGs, the second most significant source of thermal radiation on board is the set of electrical equipment operating on the spacecraft. (Figure 2.4 shows the internal arrangement of components.) From a thermal perspective, nearly all electrical systems on board perform only one function: they convert electrical energy into waste heat. The one exception is the spacecraft’s radio transmitter, converting some electrical energy into a directed beam of radio frequency energy.
Early in the missions, the RTGs produced more electrical power than what was needed on board. The main bus voltage on board was maintained at a constant value by a power supply circuit, while excess electrical power was converted into heat. The shunt regulator controlled a variable current, which flowed through the regulator itself and an external radiator plate. The external radiator plate acted as an ohmic resistor, and the power radiated by it can be easily computed from the shunt current telemetry (Figure 2.7). The shunt radiator is mounted such that its thermal radiation is primarily in a direction that is perpendicular to the spacecraft spin axis.
Some electrical heat is generated outside the spacecraft body: notable items include electrical heaters for the fuel lines that deliver fuel to the thruster cluster assemblies along the antenna rim, an electrical heater for the spacecraft battery, and some scientific instruments.
Most of the electrical power not radiated away by the spacecraft antenna or the shunt radiator is converted into heat inside the spacecraft body. As the spacecraft is very close to a thermal steady state, all the electrical heat produced internally must be radiated into space. This thermal radiation is likely to cause a measurable sunward acceleration, for several reasons. First, the spacecraft body is heavily insulated by multilayer thermal insulation; however, at the bottom of the spacecraft body is the thermal louver system designed to vent excess heat. Even in its closed state, the effective emissivity of this louver system is much higher than that of the multilayer insulation, so this is the preferred direction in which heat escapes the spacecraft. Second, the spacecraft body is situated behind the high-gain antenna; even if the thermal radiation pattern of the spacecraft body were isotropic, the back of the HGA would preferentially reflect a significant portion of this heat in the aft direction.
Figure 2.7 also indicates how the power consumption by various pieces of equipment on board can be computed from telemetry.
The thermal state of the interior of the spacecraft is characterized in a redundant manner. In addition to the thermal power of various components that can be computed from telemetry, there exist numerous temperature sensors on board, most notably among them six platform temperature sensors, which are the most likely to measure ambient temperatures (as opposed to sensors that, say, are designed to measure the temperature of a specific electrical component, such as a transistor amplifier.)
Much less is known about heat escaping the interior of the spacecraft through other routes. The spacecraft’s science instruments utilize various holes and openings in the spacecraft body in order to collect information from the environment. Design documentation prescribes that no instrument can lose more heat through the opening than its own power consumption plus 0.5 W; however, the actually heat loss per instrument is not known.
A further source of continuous heat on board is the set of radioisotope heater units (RHUs) placed at strategic locations to maintain operating temperatures. These RHUs are capsules containing a small amount of 238Pu fuel, generating 1 W of heat (Figure 2.18).
Each thruster cluster assembly housed three RHUs that prevented the freezing of thruster valves. One RHU was located at the sun sensor, while an additional RHU was placed at the magnetometer. (Some reports suggest that a 12th RHU may also have been on board one or both spacecraft.)
The thruster cluster assemblies were designed to radiate heat in a direction perpendicular to the spin axis6. There are no known asymmetries of any thermal radiation from the magnetometer assembly at the end of the magnetometer boom. Therefore, it is unlikely that thermal radiation from the RHUs in general contributed much thermal recoil force in the fore-aft direction.
The spacecraft’s propulsion system, when operating, produced significant amounts of heat; as the hydrazine monopropellant underwent a chemical reaction in the presence of a catalyst in the spacecraft’s thrusters, the thrusters and thruster cluster assemblies warmed up to several hundred degrees Centigrade. This heat was radiated into space as the assemblies cooled down to their pre-maneuver temperatures after a thruster firing event over the course of several hours.
However, the uncertainty in the magnitude of velocity change caused by the thruster event itself dwarfs any acceleration produced by the radiation of this residual heat. For this reason, heat from the propulsion system does not need to be considered when accounting for thermal recoil forces.
To complete our discussion of thermal radiation emitted by the spacecraft, we must also consider the spacecraft’s radio beam, for two reasons. First, any energy radiated by the radio beam is energy that is not converted into heat inside the spacecraft body. Second, the radio beam itself produces a recoil force that is similar in nature to the thermal recoil force.
The nominal power of the radio transmitter is 8 W, and the radio beam is highly collimated by the high gain antenna. Nevertheless, some loss occurs around the antenna fringes, and the beam itself also has a spread. Assuming that 10% of the radio beam emitted by the feedhorn misses the antenna dish, at an approximate angle of 45°, we can calculate an efficiency of 0.83 at which the antenna converts the emitted radio energy into momentum .
The actual power of the radio beam may not have been exactly 8 W. The output of the traveling wave tube oscillator that generated this microwave energy was measured on board and telemetered to the ground (Figure 2.19). Especially in the case of Pioneer 10, we note the variability of the TWT power near the end of mission. This corresponds to the drop in main bus voltage when the power available on board was no longer sufficient to maintain nominal voltages. It is unclear, therefore, if the measured drop in output power is an actual drop or a sensor artifact.
The thermal history of Pioneer 10 and 11 can be characterized accurately, and in detail, with the help of the recently recovered telemetry files and project documentation.
The telemetry data offer a redundant picture of the spacecrafts’ thermal state. On the one hand, the amount of power available on board can be computed from electrical readings. On the other hand, a number of temperature sensors on board offer a coarse temperature map of the spacecraft.
The electrical state of Pioneer 10, when the spacecraft was at 25 AU from the Sun, is shown in Figure 2.7. While this figure represents a snapshot of the spacecraft’s state at a particular moment in time, this information is available for the entire duration of both Pioneer missions.
The thermal state of the spacecraft body can be verified using the readings of six temperature sensors that were located at various points on the electronics platform: four in the main compartment, two in the adjacent compartment that housed science instruments (see locations 1 – 6 in Figure 2.9). The temporal evolution of these temperature readings agrees with expectations: outlying sensors show consistently lower temperatures, and all temperatures are dropping steadily, as the distance between the spacecraft and the Sun increases while the amount of power available on board decreases (Figures 2.20 and 2.21).
There was a large number of additional temperature sensors on board (see Appendix D). However, these temperature sensors measured the internal temperatures of on-board equipment and science instruments. One example is the hot junction temperature sensor inside the RTGs, measuring the thermocouple temperature at its hot end. These readings may be of limited use when assessing the overall thermal state of the spacecraft. Nevertheless, they are also available in the telemetry.
The thermal behavior of the spacecraft was an important concern when Pioneer 10 and 11 were designed. Going beyond a general description , much of the work that was done to ensure proper thermal behavior is documented in a review document of the Pioneer 10 and 11 thermal control subsystem . Specifically, this document provides detailed information about the thermal properties of components internal to the spacecraft, and about the anticipated maximum heat losses through various spacecraft structural and other components.
Detailed information about the SNAP-19 RTGs used on board Pioneer 10 and 11 is also available .
It has been suggested (see, e.g., ) that the material properties of these surfaces may have changed over time, introducing a time-dependent anisotropy in the spacecraft’s thermal properties. For instance, solar bleaching may affect spacecraft surfaces facing the Sun , while surfaces facing the direction of motion may be affected by the impact of interplanetary dust particles.
Note that through most of their operating lives, the high-gain antennas of the Pioneer 10 and 11 spacecraft were always pointing in the approximate direction of the Sun. Therefore, the same side of the spacecraft: notably, the interior surfaces of their HGAs and one side of each RTG, were exposed to the effects of solar light, including ultraviolet radiation, and charged particles. It is not unreasonable to assume that this continuous exposure may have altered the visible light and infrared optical properties of these surfaces, introducing in particular, a fore-aft asymmetry in the thermal radiation pattern of the RTGs. On the other hand, any such effects would be mitigated by the fact that the spacecraft were receding from the Sun very rapidly, and spent most of their operating lives at several AUs or more from the Sun, receiving only a fraction of the solar radiation that is received, for instance, by Earth-orbiting satellites of comparable age.
Another possible effect may have altered the infrared radiometric properties of spacecraft surfaces facing in the direction of motion (which, most of the time, would be surfaces facing away from the Sun.) As the spacecraft travels through the interplanetary medium at speeds in excess of 12 km/s, impact by charged particles and dust may have corroded these surfaces. Once again, this may have introduced a fore-aft anisotropy in the infrared radiometric properties of the spacecraft, most notably their RTGs.
NASA conducted several experiments to investigate the effect of space exposure on the thermal and optical properties of various materials. One investigation, called the Thermal Control Surfaces (TCS) experiment , examined the long-term effects of exposure of different materials placed on an external palette on the International Space Station. Although results of this test are quite important, the near-Earth environment is quite different from that of deep space. Concerning solar bleaching, although the spacecraft spent decades in space, most of the time was spent far away from the Sun, resulting in a low number of “equivalent Sun hours” (ESH). The cumulative exposure to solar radiation of the Sun-facing surfaces of Pioneer 10 and 11 is less than the amount of solar radiation test surfaces were exposed to during the TCS experiment, in which test surfaces were also exposed to the atomic oxygen of the upper atmosphere, which is not a consideration in the case of the Pioneer spacecraft.
The most pronounced effect on coatings in the deep space environment is due to exposure to solar ultraviolet radiation . The ESH for the Pioneer 10 and 11 spacecraft is approximately 3000 hours, most of which ( 95%) were accumulated when the spacecraft were relatively close ( 15 AU) to the Sun. The exterior surfaces of the RTGs were covered by a zirconium coating in a sodium silicate binder . For similar inorganic coatings, the most pronounced effect of prolonged solar exposure is an increase () in solar absorptance. No data suggest a noticeable change in infrared emittance.
The effects of exposure to dust and micrometeoroid impacts in the interplanetary environment do not result in significant optical damage, defined in terms of changes in solar absorptance or infrared emittance . This is confirmed by the Voyager 1 and 2 spacecraft that had unprotected camera lenses that were facing the direction of motion, yet suffered no observable optical degradation. This fact suggests that the optical effects of exposure to the interplanetary medium on the spacecraft are negligible (see discussion in [27, 392]).
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