2.2 The Pioneer spacecraft

In this section we discuss the details of the Pioneer 10 and 11 spacecraft, focusing only on those most relevant to the study of the Pioneer anomaly.

2.2.1 General characteristics

Externally, the shape of the Pioneer 10 and 11 spacecraft was dominated by the large (2.74 m diameter) high-gain antenna (HGA)3, behind which most of the spacecrafts’ instrumentation was housed in two adjoining hexagonal compartments (Figure 2.3View Image). The main compartment, in the shape of a regular hexagonal block, contained the fuel tank, electrical power supplies, and most control and navigation electronics. The adjoining compartment, shaped as an irregular hexagonal block, contained science instruments. Several openings were provided for science instrument sensors. The internal arrangement of spacecraft components is shown in Figure 2.4View Image.

The main and science compartments collectively formed the spacecraft body, which was covered by multilayer thermal insulation on all sides except part of the aft side, where a passive thermal control louver system was situated.

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Figure 2.3: A drawing of the Pioneer spacecraft. (From [292Jump To The Next Citation Point].)

For the purposes of attitude control, the entire spacecraft was designed to spin in the plane of the HGA.

Three extensible booms were attached to the main compartment, spaced at 120°. Two of these booms, both approximately 3 m long, each held two radioisotope thermoelectric generators (RTGs). This design was dictated mainly by concerns about the effects of radiation from the RTGs on the spacecrafts’ instruments, but it also had the beneficial side effect of minimizing radiative heat exchange between the RTGs and the spacecraft body. The third boom, approximately 6 m in length, held the magnetometer sensor. The length of this boom ensured that the sensor was not responding to magnetic fields originating in the spacecraft itself.

The total mass of Pioneer 10 and 11 was approximately 260 kg at the time of launch, of which approximately 30 kg was propellant and pressurant. The masses of the spacecraft slowly varied throughout their missions primarily due to propellant usage (for details, see Section 2.3.2).

The propulsion system was designed to perform three types of maneuvers: spin/despin (setting the initial spin rate shortly after launch), precession (to keep the HGA pointing towards the Earth, and also to orient the spacecraft during orbit correction maneuvers) and velocity changes. Of the two spacecraft, Pioneer 11 used more of its propellant, in the course of velocity correction maneuvers that were used to adjust the spacecraft’s trajectory for its eventual encounter with Saturn.

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Figure 2.4: Pioneer 10 and 11 internal equipment arrangement. (From [292Jump To The Next Citation Point].)

2.2.2 Science instruments

The Pioneer spacecraft carried an identical set of 11 science instruments, with a 12th instrument present only on Pioneer 11, namely:

JPL Helium Vector Magnetometer
ARC Plasma Analyzer
U/Chicago Charged Particle Experiment
U/Iowa Geiger Tube Telescope
GSFC Cosmic Ray Telescope
UCSD Trapped Radiation Detector
UCS Ultraviolet Photometer
U/Arizona Imaging Photopolarimeter
CIT Jovian Infrared Radiometer
GE Asteroid/Meteoroid Detector
LaRC Meteoroid Detector
Flux-Gate Magnetometer (Pioneer 11 only)

The power system of the Pioneer spacecraft was designed to ensure that a successful encounter with Jupiter (with all science instruments operating) can be carried out with only three (out of four) functioning radioisotope thermoelectric generators. Instruments could be commanded on or off by ground control; late in the extended mission, when sufficient power to operate all instruments simultaneously was no longer available, a power sharing plan was implemented to ensure that the power demand on board would not exceed available power levels.

At the end of its mission, only one instrument on board Pioneer 10 remained in operation; the University of Iowa Geiger Tube Telescope. An attempt was made to power down this instrument, in order to improve the available power margin. It is not known if this command was received or executed by the spacecraft.

2.2.3 Radioisotope thermoelectric generators (RTGs)

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Figure 2.5: The SNAP-19 RTGs used on Pioneer 10 and 11 (from [350Jump To The Next Citation Point]). Note the enlarged fin structure. Dimensions are in inches (1” = 2.54 cm).

The primary electrical source on board Pioneer 10 and 11 was a set of four radioisotope thermoelectric generators [1135350Jump To The Next Citation Point38439]. Each of these RTGs contains 18 238Pu capsules, approximately two inches (5.08 cm) in diameter and 0.2 inches (0.51 cm) thick. The total thermal power of each RTG, when freshly fueled, was ∼ 650 W. Each RTG contains two sets of 45 bimetallic thermocouples, connected in series. The thermocouples initially operated at ∼ 6% efficiency; the nominal RTG output is ∼ 4 V, 10 A. The total available power at launch on board each spacecraft was ∼ 160 W (Figure 2.14View Image).

Excess heat from each RTG is radiated into space by a set of six heat radiating fins. The fins provide the necessary radiating area to ensure that a sufficient temperature differential is present on the thermocouples, for efficient operation.

The thermal power of the RTGs is a function of the total power produced by the 238Pu fuel, and the amount of power removed in the form of electrical energy by the thermocouples. The half-life of 238Pu is 87.74 years. The efficiency of the thermocouples decreased over the years as a result of the decreasing temperature differential between the hot and cold ends, and also as a result of aging. At the time of last transmission, each RTG on board Pioneer 10 produced less than 15 W of electrical power.

The actual shape of the RTGs is shown in Figure 2.5View Image. It is important to note that the RTGs on board Pioneer 10 and 11 were built with the large fins that are depicted in this figure. These enlarged fins are not shown in many drawings and photographs (including Figure 2.3View Image).

2.2.4 The electrical subsystem

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Figure 2.6: Overview of the Pioneer 10 and 11 electrical subsystem (from [292Jump To The Next Citation Point]).

Electrical power from the RTGs reached the spacecraft body via a set of ribbon cables. There, raw power from the RTGs was fed into a series of electric power supplies that produced 28 VDC for the spacecraft’s main bus, and also other voltages on secondary power buses.

Electrical power consumption on board the spacecraft was regulated to ensure a constant voltage on the various power buses on the one hand, and an optimal current draw from the RTGs on the other hand. The electrical power subsystem was designed such that the spacecraft could perform its primary mission, namely close-up observations of the planet Jupiter approximately 21 months after launch, using only three RTGs, while operating a full compliment of science instruments. Consequently, in the early years of their mission, significant amounts of excess electrical power were available on both spacecraft. The power regulation circuitry diverted this excess power to a shunt circuit, which dissipated some excess power internally, while routing the remaining excess power to an externally mounted shunt radiator.

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Figure 2.7: Pioneer 10 power budget on July 25, 1981, taken as an example. Power readings that were obtained from spacecraft telemetry are indicated by the telemetry word in the form Cnnn. The discrepancy between generated power and power consumption is due to rounding errors and uncertainties in the nominal vs. actual power consumption of various subsystems.

The total power available on board at the time of launch was in excess of ∼ 160 W. This figure decreased steadily throughout the missions, obeying an approximate negative exponential law. The actual amount of power available was a function of the decay of the radioisotope fuel, the decreasing temperature differential between the hot and cold ends of thermoelectric elements, and degradation of the elements themselves. At the end of its mission, the power available on board Pioneer 10 was less than 60 W. (Indeed, a drop in the main bus voltage is the most likely reason that Pioneer 10 eventually fell silent, as the reduced voltage was no longer sufficient to operate the spacecraft’s transmitter.)

The RTGs generated electrical power at ∼ 4 VDC (Figure 2.6View Image). Power output from each RTG was fed to a separate inverter circuit, producing 61 VAC (peak-to-peak) at ∼ 2.5 kHz. Output from the four inverters was combined and fed to the Power Control Unit (PCU), which generated the 28 VDC main bus voltage, managed the on-board battery, controlled the dissipation of excess power via a shunt circuit, and also provided power to the Central Transformer Rectifier (CTRF) component, which, in turn, supplied power at various voltages (e.g., ± 16 VDC, ± 12 VDC, +5 VDC) to other subsystems and instruments.

The power budget at any given time was a function of available power vs. spacecraft load. For instance, Figure 2.7View Image shows Pioneer 10’s power budget on July 25, 1981.

The on-board battery was designed to help with transient peak loads that temporarily exceeded the capabilities of the RTGs. The battery was composed of eight silver-cadmium cells, each of which had a capacity of 5 Ah and was equipped with an individual charge/discharge bypass circuitry.

2.2.5 Propulsion and attitude control

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Figure 2.8: An overview of the Pioneer 10 and 11 propulsion subsystem (from [292Jump To The Next Citation Point]).

After separation from their respective launch vehicles, the orbits of Pioneer 10 and 11 were determined by the laws of celestial mechanics. The spacecraft had only a small amount of fuel on board, used by their propulsion system designed to control the spacecraft’s spin and orientation, and execute minor course correction maneuvers.

The propulsion system consisted of three thruster cluster assemblies (see Figure 2.8View Image), each comprising two 1 lb (∼ 4.5 N) thrusters. All thruster cluster assemblies were mounted along the rim of the HGA. One pair of clusters was oriented tangentially along the antenna perimeter, and its two thrusters were intended to be used to increase or decrease the spin rate of the spacecraft. The remaining two pairs were oriented perpendicular to the antenna plane, on opposite sides of the antenna. These two thruster cluster assemblies were used in pairs. If two thrusters pointing in the same direction were fired simultaneously, this resulted in a net change in the spacecraft’s velocity in a direction perpendicular to the antenna plane. If two clusters were fired in the opposite direction, this caused the spacecraft’s spin axis to precess. This latter type of maneuver was used, in particular, to maintain an Earth-pointing orientation of the HGA to ensure good reception of radio signals.

The thrusters were labeled VPT (velocity and precession thruster) and SCT (spin control thruster.) VPT 1 and VPT 3 were oriented in the same direction as the HGA (the +z direction), while VPT 2 and VPT 4 were oriented in the opposite direction.

The propulsion system utilized hydrazine (N2H4) monopropellant fuel, of which ∼ 27 kg was available on board, in a 38 liter tank that was pressurized with N2. The propellant and pressurant were separated by a flexible membrane, which prevented the mixing of the liquid propellant and gaseous pressurant in the weightless environment of space. The fuel tank was located at the center of the spacecraft, and was heated by the spacecraft’s electrical equipment. Fuel lines leading to the thruster cluster assemblies were heated electrically, while the thruster cluster assemblies were equipped with small (1 W) radioisotope heating units (RHUs) containing 238Pu fuel.

The capabilities of the propulsion system are summarized in Table 2.2.

Table 2.2: Capabilities of the Pioneer 10 and 11 propulsion system.
Maneuver Thrusters Max. capability
Despin SCT 1 or 2 58 rpm
Spin control SCT 1 or 2 14 rpm
Precession VPT 1 and 4 or VPT 2 and 3 1250°
Delta-v VPT 1 and 3 or VPT 2 and 4 250 m/s

The spacecraft’s Earth-pointing attitude was maintained as the spacecraft were spinning in the plane of the HGA, at a nominal rate of 4.8 revolutions per minute (rpm). The propulsion system had the capability to adjust the spin rate of the spacecraft, and to precess the spin axis, in order to correct for orientation errors, and to ensure that the spacecraft followed the Earth’s position in the sky as seen from on board.

The spin axis perpendicular to the plane of the HGA is one of the spacecraft’s principal axis of inertia. A wobble damper mechanism [292Jump To The Next Citation Point] dampened rotations around any axis other than this principal axis of inertia, ensuring a stable attitude even after a precession maneuver.

2.2.6 Navigation

The Pioneer 10 and 11 spacecraft relied on standard methods of deep space navigation [235Jump To The Next Citation Point240Jump To The Next Citation Point] (see Section 4). The spacecraft’s position was determined using the spacecraft’s radio signal and the laws of celestial mechanics. The radio signal offered a precision Doppler observable, from which the spacecraft’s velocity relative to an Earth station along the line-of-sight could be computed. Repeated observations and knowledge of the spacecraft’s prior trajectory were sufficient to obtain highly accurate solutions of the spacecraft’s orbit.

The orientation of the spacecraft was estimated from the quality of the radio communication link (i.e., the spacecraft had to be approximately Earth-pointing in order for the Earth to fall within the HGA radiation pattern.) The rate and phase of the spacecraft’s rotation was established by a redundant pair of sun sensors and a star sensor on board. These sensors (selectable by ground command) provided a roll reference pulse that was used for navigation purposes (as explained in the next paragraph) as well as by on-board science instruments. The time between two subsequent roll reference pulses was measured and telemetered to the ground.

The spacecraft also had minimal autonomous (closed loop) navigation capability, designed to make it possible for the spacecraft to restore its orientation by “homing in” on an Earth-based signal. The maneuver, called a conical scan (CONSCAN) maneuver, utilized a piston mechanism [3] with electrically heated freon gas that displaced the feed horn located at the focal point of the high-gain antenna. Unless the Earth was exactly on the HGA centerline, this introduced a sinusoidal modulation in the amplitude of the signal received from the Earth. A simple integrator circuit, utilizing this sinusoidal modulation and the roll reference pulse, triggered firings of the precession thrusters, adjusting the spacecraft’s axis of rotation until it coincided with the direction of the Earth.

Frequently, instead of CONSCAN maneuvers, “open loop” attitude correction maneuvers were used, calculated to ensure that after the maneuver, the spacecraft was oriented to “lead” the Earth, allowing the Earth to move through the antenna pattern subsequently. This reduced the frequency of attitude correction maneuvers; further, open loop maneuvers generally consumed less propellant than autonomous CONSCAN maneuvers.

Early in the mission, attitude correction maneuvers had to be executed regularly, due to the combined motion of the Earth and the spacecraft. Late in the extended mission, only two attitude correction maneuvers were needed annually, to compensate for the Earth’s motion around the Sun, and for the spacecraft’s “sideways” motion along its hyperbolic escape trajectory.

2.2.7 Communication system

The spacecraft maintained its communication link with the Earth using a set of S-band transmitters and receivers on board, in combination with three antennae.

The main communication antenna of the spacecraft was the 2.74 m diameter high-gain antenna. The antenna’s narrow beamwidth (3.3° downlink, 3.5° uplink) ensured an effective radiated power of 70 dBm, allowing communication with the spacecraft over interplanetary distances. (The original mission design anticipated signal loss some time after Jupiter encounter, but still within the orbit of Saturn; increases in the sensitivity of Earth stations allowed communication with Pioneer 10 up until 2003, when the spacecraft was over 70 AU from the Earth.)

Mounted along the centerline of the HGA was the horn of a medium-gain antenna (MGA). On the opposite (–z) side of the spacecraft, at the bottom of the main compartment was mounted a third, low-gain omnidirectional antenna (LGA). This antenna was used during the initial mission phases, before the HGA was oriented towards the Earth.

The spacecraft had two receivers and two transmitters on board, switchable by ground command. While one receiver was connected to the HGA, the other was connected to the MGA/LGA. The sensitivity of the receiver was –149 dBm; at the time of the last transmission, the spacecraft detected the Earth station’s signal at a strength of –131.7 dBm.

The spacecraft utilized two traveling wave tube (TWT) transmitters for microwave transmission. The TWTs were selectable by ground command; it was possible to power off both TWTs to conserve power (such as when CONSCAN maneuvers were performed late in the mission, when the available electrical power on board was no longer sufficient to operate a TWT transmitter and the feed movement mechanism simultaneously.)

The radio systems operated in the S-band, utilizing a frequency of ∼ 2.1 GHz for uplink, and ∼ 2.3 GHz for downlink. The transmitter frequency was synthesized on board by an independent oscillator. However, the spacecraft’s radio system could also operate in a coherent mode: in this mode, the downlink signal’s carrier frequency was phase-coherently synchronized to the uplink frequency, at the exact frequency ratio of 240/221. In this mode, the precision and stability of the downlink signal’s carrier frequency was not limited by the equipment on board. This mode allowed precision Doppler frequency measurements with millihertz accuracy.

The main function of the spacecraft’s communication system was to provide two-way data communication between the ground and the spacecraft. Data communication was performed at a rate of 16 – 2048 bits per second (bps). Communication from the ground consisted of commands that were decoded by the spacecraft’s radio communication subsystem. Communication to the ground consisted of measurement results from the spacecraft’s suite of science instruments, and engineering telemetry.

2.2.8 Telemetry data and its interpretation

Communication between the spacecraft and a DSN antenna took place using a variety of data formats; the format selected depended on the type of experiment that the spacecraft was ordered to perform, but typically the format gave precedence to science results over telemetry. However, telemetry information was continuously transmitted to the Earth at all times, using a small portion of the available data bandwidth.

The telemetry data stream was assembled on board by the digital telemetry unit (DTU), which comprised some ∼ 800 TTL4 integrated circuits, and was also equipped with 49,152 bits of ferrite core data memory. A total of 10 science data formats (of which 5 were utilized) and 4 engineering data formats, yielding a total number of 18 different valid format combinations, was selectable by ground command. The DTU could operate in three modes: realtime (passing through science measurements as received from instruments), store (storing measurements in the on-board memory) and readout (transmitting measurements previously stored in on-board memory.)

When one of the engineering data formats was selected, the spacecraft transmitted only engineering telemetry. These formats were utilized, for instance, during maneuvers or spacecraft troubleshooting.

Most of the time, a science data format was used, in which case most of the bits in the telemetry data stream contained science data. A small portion, called the subcommutator, was reserved for engineering data; this part of the telemetry record cycled through all telemetry data words in sequence.

The telemetry record size was 192 bits, divided into 36 6-bit words. Depending on the telemetry format chosen, either all 36 6-bit words contained engineering telemetry, or a single engineering telemetry word was transmitted in every (or every second) telemetry record.

There were 128 distinct engineering telemetry words. Depending on the science format used, the subcommutator was present either in every transmitted record or every second record. Therefore, it may have taken as many as 256 telemetry records before a particular engineering word was transmitted. At the lowest data rate of 16 bits per second, this meant that any given parameter was telemetered to the Earth once every 51.2 minutes.

An additional 64 telemetry words were used to transmit engineering information from science instruments. These parameters were only transmitted in the subcommutator, and at the lowest available data rate, a particular science instrument telemetry word was repeated every 25.6 minutes.

The 128 engineering telemetry words were organized into 4 groups of 32 words each; it is customary to denote them using the notation Cmnn, where m = 1...4 is the group number, and nn = 01...32 is the telemetry word. Similarly, science instrument telemetry words were labeled Emnn, with m = 1...2 and nn = 01...32.

When a science or engineering telemetry word was used to convey the reading from an analog (temperature, voltage, current, pressure, etc.) sensor, the reading was digitized with 6-bit resolution [146402Jump To The Next Citation Point167], and the resulting 6-bit data word was transmitted. Associated with each telemetry word representing an analog measurement was a set of calibration coefficients that formed a 5th-order calibration polynomial. For each calibration polynomial, a range was also defined that established valid readings that could be decoded by that polynomial.

Appendix D lists selected engineering and science telemetry data words that may be relevant to the analysis of the Pioneer anomaly.

2.2.9 Thermal subsystem

The Pioneer spacecraft were equipped with a thermal control system comprising a variety of active and passive thermal control devices. The purpose of these devices was to maintain the required operating temperatures for all vital subsystems of the spacecraft.

The main spacecraft body was covered by multilayer insulating blankets [364365]. These blankets were designed to retain heat within the spacecraft when it was situated in deep space, far from the Sun.

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Figure 2.9: Location of thermal sensors in the instrument compartment of the Pioneer 10 and 11 spacecraft (from [292Jump To The Next Citation Point]). Platform temperature sensors are mounted at locations 1 to 6. Some locations (i.e., end of RTG booms, propellant tank interior, etc.) not shown.

To prevent overheating of the spacecraft interior near the Sun, a thermal louver system [292Jump To The Next Citation Point79299385Jump To The Next Citation Point] was utilized. The louvers were located at the bottom of the spacecraft, organized in a circular pattern, with additional louvers on the science compartment (Figure 2.10View Image). These louvers were actuated by bimetallic springs that were thermally (radiatively) coupled to the main electronics platform behind the louvers. The louvers were designed to be fully open when the platform temperature exceeded 90° F, and fully close when the temperature fell below 40° F (Figure 2.11View Image).

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Figure 2.10: The Pioneer 10 and 11 thermal control louver system, as seen from the aft (–z) direction (from [292Jump To The Next Citation Point]).
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Figure 2.11: Louver blade angle as a function of platform temperature (from [385Jump To The Next Citation Point]). Temperatures in ° F ([° C] = ([° F]–32) × 5/9).
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Figure 2.12: Louver structure heat loss as a function of platform temperature (from [385Jump To The Next Citation Point]). Temperatures in ° F ([° C] = ([° F]–32) × 5/9).
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Figure 2.13: Louver assembly performance (from [385Jump To The Next Citation Point]). Temperatures in ° F ([° C] = ([° F]–32) × 5/9).

The louver system emits heat in two ways: through structural components (Figure 2.12View Image), and through the louver blade assemblies (Figure 2.13View Image). The total heat emitted by the louver system is the sum of the heat emitted via these two mechanisms.

The fuel lines extending from the spacecraft body to the thruster cluster assemblies located along the rim of the HGA were insulated by multilayer thermal blankets and were kept warm by electric heater lines. The thruster cluster assemblies contained radioisotope heater units (RHUs), designed to prevent the propellant from freezing. Additional RHUs heated the star sensor and the magnetometer assembly at the end of the magnetometer boom.

The spacecraft’s battery was mounted on the outside of the main spacecraft body, and heated by an electrical heater.

Most heat produced by electrical equipment on board was released within the insulated interior of the spacecraft. The requirement for science instruments was to leak no more than the instrument’s own power consumption plus 0.5 W to space.

Some excess electrical power was radiated away as heat by an externally mounted shunt radiator plate, which formed part of the spacecraft’s electrical power subsystem.

Most heat on board was produced by the four radioisotope thermoelectric generators. Electrical power in these generators was produced by bimetallic thermocouples that relied on the temperature difference between their hot and cold ends for power generation. Therefore, it was essential that the cold ends of the thermocouples were connected to RTG radiator fins that radiated heat into space with high efficiency.

Temperature readings from many locations throughout the spacecraft, including temperatures at six key locations on the main electronics platform, were telemetered to the ground (Figure 2.9View Image).

With the exception of the louver system, the two adjacent hexagonal parts of the spacecraft body are covered by multilayer insulation. There is no insulation between the main and science compartments of the spacecraft body.

The surface materials and paints used to cover most major exterior surfaces are documented [292Jump To The Next Citation Point61Jump To The Next Citation Point114144206385Jump To The Next Citation Point]. In particular, the surfaces of the spacecraft body, HGA, and RTGs are well described, along with the thermal control louver system in terms of solar absorptance and infrared emittance (Table 2.3). Solar absorptance is characterized by a dimensionless number (usually denoted by α) between 0 and 1 representing the efficiency with which a particular material absorbs the radiant energy of the Sun when compared to an ideal black body. Infrared emittance is similarly characterized by a dimensionless number (usually dented by 𝜖) between 0 and 1 that represents the efficiency with which a material radiates heat at lower (typically, room) temperatures as compared to an ideal black body.

Table 2.3: Radiometric properties of Pioneer 10 and 11 major exterior surfaces (at launch): solar absorptance (α) and infrared emittance (𝜖).
Surface Area (m2) α 𝜖
HGA front (white paint) 5.91 0.21 0.85
HGA rear (bare aluminum) 5.91 0.17 0.04
Spacecraft body, front (2-mil aluminized Mylar) 1.55 0.17 0.70
Spacecraft body, rear (2-mil aluminized Kapton) 1.19 0.40 0.70
Spacecraft body, RTG sides (2-mil aluminized Kapton) 0.65 0.40 0.70
Spacecraft body, other sides (2-mil aluminized Mylar) 1.21 0.17 0.70
RTG fin surfaces (white paint)   0.20 0.83
Louver blades, closed (bare aluminum) 0.36 0.17 0.04

The exterior surfaces of the Pioneer 10 and 11 spacecraft are covered by a variety of materials and paints. Changes in the spacecraft properties can affect/induce forces that are both of on-board and of external origin. Therefore, it is of great importance to establish the extent to which any of the spacecraft’s properties might have been changing over time.

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