The Global Precipitation Measurement ( GPM ) Spacecraft Power System Design and Orbital Performance

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA). It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The spacecraft is in a circular 400 Km altitude, 65 degrees inclination nadir pointing orbit with a three year basic mission life. The solar array consists of two sun tracking wings with cable wraps. The panels are populated with triple junction cells of nominal 29.5% efficiency. One axis is canted by 520 to provide power to the spacecraft at high beta angles. The power system is a Direct Energy Transfer (DET) system designed to support 1950 Watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s x 84p batteries operated in parallel as a single battery. The paper describes the power system design details, its performance to date and the lithium ion battery model that was developed for use in the energy balance analysis and is being used to predict the on-orbit health of the battery. INTRODUCTION The Global Precipitation Measurement (GPM) spacecraft, shown in Fig. 1 is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014 from Tanegashima, Japan for a three year basic life and five years extended mission life. GPM succeeds the Tropical Rainfall Measuring Mission (TRMM) spacecraft, which was also a joint mission between National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA) to study rainfall for weather and climate research that operated in a 400Km 65 ̊ inclination orbit. The TRMM satellite mission ended on April 15, 2015 after 17 years of operation. Since launch, GPM has been providing accurate, frequent global detailed measurements of precipitation to assist researchers in studying global climate, and improving the forecasting of extreme weather events. The spacecraft has two instruments; the Dual-frequency Precipitation Radars (DPR) and the GPM Microwave Imager (GMI). The spacecraft development was funded jointly by NASA and JAXA. NASA/Goddard Space Flight Center (GSFC) was responsible for the spacecraft, the GMI instrument, performed the Integration and Test (I&T) and is responsible for the Mission Operations from GSFC. JAXA built the DPR instrument, provided the launch vehicle and performed the launch services. The spacecraft is in a circular 400 Km altitude, 65 degree inclination orbit. The GPM orbit and inclination were selected to provide full global coverage updated in a 24 hour period. The beta angle varies during the year approximately -90 and +90 degrees. The orbit period is 93 minutes. The maximum eclipse duration is about 37 minutes. Fig. 2 shows the beta angle and eclipse duration variation during the year. The spacecraft is nadir pointing in the +Z direction and flies in the X direction. To maintain the DPR instrument pointed away from the sun with orbit precession during the year, 180 degrees yaw manoeuvres are performed at around beta zero. The solar array consists of two sun tracking wings with cable wraps. Each wing consists of four panels. One axis is canted by 520 to provide power to the spacecraft at high beta angles. Figure 1. GPM Spacecraft On-Orbit Configuration Figure 2. Eclipse time (red) and beta angle (black) variations during a year https://ntrs.nasa.gov/search.jsp?R=2016001202


INTRODUCTION
The Global Precipitation Measurement (GPM) spacecraft, shown in Fig. 1 is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014 from Tanegashima, Japan for a three year basic life and five years extended mission life.GPM succeeds the Tropical Rainfall Measuring Mission (TRMM) spacecraft, which was also a joint mission between National Aeronautics and Space Administration (NASA) and Japan Aerospace Exploration Agency (JAXA) to study rainfall for weather and climate research that operated in a 400Km 65 ࡈ inclination orbit.The TRMM satellite mission ended on April 15, 2015 after 17 years of operation.Since launch, GPM has been providing accurate, frequent global detailed measurements of precipitation to assist researchers in studying global climate, and improving the forecasting of extreme weather events.The spacecraft has two instruments; the Dual-frequency Precipitation Radars (DPR) and the GPM Microwave Imager (GMI).The spacecraft development was funded jointly by NASA and JAXA.NASA/Goddard Space Flight Center (GSFC) was responsible for the spacecraft, the GMI instrument, performed the Integration and Test (I&T) and is responsible for the Mission Operations from GSFC.JAXA built the DPR instrument, provided the launch vehicle and performed the launch services.The spacecraft is in a circular 400 Km altitude, 65 degree inclination orbit.The GPM orbit and inclination were selected to provide full global coverage updated in a 24 hour period.The beta angle varies during the year approximately -90 and +90 degrees.The orbit period is 93 minutes.The maximum eclipse duration is about 37 minutes.Fig. 2 shows the beta angle and eclipse duration variation during the year.The spacecraft is nadir pointing in the +Z direction and flies in the X direction.To maintain the DPR instrument pointed away from the sun with orbit precession during the year, 180 degrees yaw manoeuvres are performed at around beta zero.The solar array consists of two sun tracking wings with cable wraps.Each wing consists of four panels.One axis is canted by 52ͼ to provide power to the spacecraft at high beta angles.The Electrical Power System (EPS) consists of: The Power System Electronics (PSE), three Sony 18650HC based lithium ion batteries and two solar array wings.The power system topology and PSE hardware elements have heritage to Solar Dynamics Observatory (SDO), Lunar Reconnaissance Orbiter (LRO) and other recent NASA/GSFC spacecraft.The power system is designed to support 1950 watts orbit average power.The solar array wings are divided into segments.Segments from each wing are connected to the power bus in the PSE through isolation diodes in the "Segment Module".A transistor connected from the diode anode to power return is used to shunt the circuit from the power bus when the solar array segment power is not needed to perform "coarse" power control.There are 24 switched segments.The fine solar array control is performed by redundant boost topology Pulse Width Modulated (PWM) converter in the Solar Array Module (SAM).The solar array strings connected to each of the redundant SAM PWM modules are from strings of each wing.The PWM module strings of the flat +Y wing experience extensive shadowing caused by the spacecraft body as well as the spinning GMI instrument antenna dish.The PWM solar cell strings from the +Y wing are placed on the outermost panel that is the last to be shadowed with increased beta angle.The PWM circuit solar array strings from the canted -Y wing, which is always in the Sun, are located on the second panel.The PWM converters operate at a fixed 125 KHz.Only one PWM controller is in control at a time.

Power System Electronics (PSE)
The PSE contains the circuits needed to control the solar array power, charge the batteries as well as the electronics for power distribution and current monitoring of spacecraft loads.Fig. 4 shows the PSE simplified block diagram indicating the power services and their current ratings.The PSE is in one box that is internally redundant.Both primary and redundant sides A&B are powered simultaneously during normal operations.The PSE contains the electronics to perform regulation of power bus, battery charging and solar array power control with the PWM SAM and digital segments control modules, monitoring battery voltages, currents and temperatures.It contains the circuitry for spacecraft power distribution using over-current protected switched services using hybrid solid state resettable circuit breakers.The PSE also contains fuse protected unswitched power services.The fuses are placed in accessible Fuse Plugs on the PSE.The PSE interfaces to the spacecraft Communication and Data Handling (C&DH) system on Primary and Redundant Mil-STD-1553 interfaces.In addition, there are Special Hardline interface commands and telemetry lines to the C&DH for recovery from Mil-STD-1553 bus and internal PSE failures.The PSE contains the Umbilical and ground Direct Power Conditioner (DPC) interfaces.The spacecraft Central Ground Point is in the PSE.The PSE controls the battery charging with current limited selectable levels until the battery voltage reaches a ground selectable voltage limit.The voltage is then held constant while the charge current tapers toward zero.The current and voltage limits are set with 12 bit Digital to Analog Converters (DAC)s.The battery current control loop also allows controlled constant current discharge at selectable levels.This feature was used during ground testing to simplify and accelerate battery discharge.In flight, it is used to perform special tests such as battery internal resistance characterization during full sun periods.In addition to the battery current and voltage control loops, a separate power bus overvoltage control loop is incorporated in the design to limit the power bus voltage to safe levels.This protects the spacecraft from inadvertent removal of the batteries from the power bus (particularity during ground testing).This feature also protects the bus from over-voltage conditions in the event of a failure in the voltage or current control loops.The overvoltage control is performed by a hysteresis controller that switches ON / OFF groups of eight solar array (SA) segment shunting transistors.During the overvoltage control loop operation, the bus voltage oscillates very slowly as the battery charges and discharges with removal or addition of the group of segments depending on loads and Solar Array Simulator (SAS) power settings.Four overvoltage limits are available.The PSE is housed in a single machined aluminum alloy box.The overall enclosure dimensions are 245 mm x 785 mm x 319 mm (HxLxW), the mass is 47.3 Kg.The PSE uses parts selected to a radiation dosage of 10 KRad total ionizing dose (TID) with 2.54 mm Al shielding, but many parts were 100 KRad parts in order to meet the single event effects requirements.

Batteries
The GPM batteries were built by ABSL, U.S.A. in Longmont, Colorado, Using SONY 18650HC cells.The battery set consists of three 8s x 84p batteries operated in parallel as a single battery.The nameplate End-of-Life (EOL) battery capacity is 80 Ampere-Hour (Ahr) when tested to an end-of-discharge voltage of 24 V (3 V/cell).I&T batteries, identical to flight units but with cells of a different lot, were used during spacecraft I&T.The flight batteries were installed at the launch site, as is typical in GSFC battery management plans for GPM class spacecraft.Each battery consists of seven 8s x 12p blocks arranged in two decks; three blocks at lower deck and 4 blocks on upper deck.One block volume at the lower deck, houses the two parallel battery disconnect 100 A Hartman relays, the connectors wiring and power bus pc boards with voltage sense resistors and relay drive decoupling diodes.Each battery has six 2 K Platinum Resistor Thermometers (PRT) mounted on selected cells at the predicted hottest and coldest locations on six of the seven blocks to monitor battery temperature during flight.In DGGLWLRQ WKHUH DUH WKUHH .WKHUPLVWRUV PRXQWHG DW predicted hottest and coldest cells for ground I&T monitoring.The three batteries are mounted on a single baseplate with heat pipes to maintain low temperature gradients between the three battery cells.Figure 5 shows

Solar Array
The two articulated solar array wings rotate independently around the Y axis.Cable wraps are used to transfer the power from the rotating wings to the spacecraft.The cable wraps are rewound during the eclipse period.Fig. 6 shows the two deployed wings.The solar cell laydown was performed by EMCORE in Albuquerque, NM, USA using their ZTJ cell with nominal 29.5% efficiency.The GSFC provided panel substrates were aluminium honeycomb core with composite facesheets.The cell side was co-cured with 50 μm Kapton insulator sheets.The backside of the panels was painted with A-276 white thermal paint.Protection of the Kapton insulator and the silver plating on cell interconnect from Atomic Oxygen (AO) was required.

Spacecraft Instrument Operation Transients Effect on Batteries Investigation
At initial DPR instrument integration with the spacecraft at GSFC, unexpected high current transients were noticed on the DPR power lines and the batteries when the DPR was operated.A typical scope current wave shape in one the three batteries is shown in Fig. 7. Investigation revealed that the DPR phased array radar generates cyclical high rate large current transients on the spacecraft power bus.The interaction of the DPR high current transients with the limited bandwidth PSE operation lead to the battery transient current profile observed.The pattern repeated at 700 milliseconds with the low amplitude component frequency of the transients being around 75 Hz.Small discharge/charge battery cycling will occur during full sun periods when the batteries are fully charged with zero charge current and during short eclipse orbits where the charge current tapers to very low values.The small discharge/charge cycling impact on the battery was a concern and was investigated [1].
An accelerated life test was conducted using two GPM I&T batteries.One battery was used as the test battery, the second as the control.A worst case piecewise linear profile simulating the observed wave shape consisted of four ±3 A square wave pulses, followed by ±0.5 A square wave for about 30 cycles.The pattern was repeated at 10 Hz rate.The two batteries were maintained at room temperature during the life test.Periodic capacity tests were conducted.Tests over a 15 month period did not indicate measurable difference between the reference and test batteries.The test profile was changed to a ±2.5 A 100 Hz test.Capacity test after six month with the new profile did not indicate noticeable difference in degradation due to the micro discharge/charge cycling either indicating that there should be negligible impact on GPM batteries life due to the DPR transients.

Figure 7. DPR instrument induced currents in each GPM battery Power System Orbital Operation
The actual orbital load power of the spacecraft components particularly the instruments and spacecraft heater power was considerably lower than the 1950 W orbit average design specification.Orbital data indicated a power demand of around 1250 W at low beta and around 1350 W at high beta angles with a corresponding maximum eclipse battery Depth-of-Discharge (DoD) of around 10%.To take advantage of the low spacecraft power with its resultant low battery DoD, the operational charge voltage was lowered to reduce the stresses in the battery; the battery operation was at 33.6 V at launch and during the very early period after launch, It was lowered to 33.1V (around 90% State-of Charge (SOC)).It was further reduced after about one year to 32.9 V (around 85% SOC).Since September 2015, it is operating at a voltage limit of 32.7 V corresponding to about 4.1 V/cell (around 80% SOC).The batteries still store ample energy for emergencies at this lower SOC.On board software for solar panel sun-pointing control was modified to reduce the drag caused by the large solar panels during flight by feathering the two wings alternatively depending on beta angle variations; the +Y wing is made to track between beta angles 0 to ± 20 while -Y wing is feathered with its panel edges in the velocity vector direction, while for beta angles between ±20 to ±55 degrees, the +Y is feathered with its edge in the velocity vector direction while +Y wing is tracking the sun, then for beta angles above ±55 degrees, both wings are feathered.Feathering the wings minimize the projected area in the velocity vector directions and lowers substantially the drag on the spacecraft which reduces the number of propulsion burns required to maintain the spacecraft in the desired altitude.Feathering also reduces the power generated by the solar array that must be processed by the PSE leading to several degrees reduction in PWM controller transistor and diode temperatures.In addition, feathering require fewer solar array drive movements thus less stress and wear on solar array drive system.The reduction in propulsion maneuvers for altitude adjustment has resulted in considerable savings in propulsion fuel to date.It will allow further mission extensions.Figure 13 shows the orbit average temperature variations of battery-1, the plots for the other two batteries are identical.

Battery Modelling and Power System Energy Balance
A detailed GPM energy balance program was developed which included the lithium ion battery model, the solar array model, Sun to panels orbital angle variations, solar array temperature, solar array shadowing effects, power system solar array/battery control electronics and load models.The energy balance program was used as part of the solar array sizing analysis, verification for EOL worst case conditions, predicting the battery and power system performance for various mission operation scenarios and the impact on the available spacecraft power of various fault scenarios.It is presently used for battery health    Battery on-orbit internal resistance characterization is also being performed during full sun periods by utilizing the PSE capability for setting the charge current to a negative value.After stabilizing the taper current to zero at a selected operating voltage, the batteries are discharged at constant rates starting from same SOC.The internal resistance is derived as delta V/delta I.The Rin characterization were conducted during recent two full sun periods about 3 month apart.The differences were within measurement uncertainties during the period.

Figure 18. Battery voltage flight and model results with lunar shadowing profile Summary and Conclusions
The GPM spacecraft has been providing important science data for global weather monitoring and predictions.The power system performance has been excellent.There is ample solar array power and the battery DoD is low.Operational changes implemented reduced the propellant use and the stresses in the battery and electronic components.To date there have been two -Y wing panels PRT temperature sensors that have been intermittent.These two PRTs provide data as telemetry information that are not critical for power system operation.The power degradation in the battery and solar array have been less than predictions.The power system is expected to support the mission for substantially longer period than the baseline extended mission plan.

Figure 4 .
Figure 4. PSE simplified block diagram photo of one GPM battery.A single battery mass and dimensions is 41.2 Kg and 676 mm x 306 mm x 166 mm.

Figure 6 .
Figure 6.Solar array deployed wings.-Y top, +Y bottom AO is prevalent at the 400 km altitude orbit and was predicted to be 1.464 x 10 22 atoms/cm 2 during the mission.AO erodes the silver converting it into silver oxide and destroys the Kapton insulator exposing the conductive composite facesheet.Prior to cell laydown, GSFC thermal coating personnel sprayed a few microns of low viscosity room temperature vulcanization (RTV) Fig. 8 show the observed variation of Voc and Isc to date adjusted for temperature, solar distance and angle.The ࡈ &, one Sun AM0, normal sun incidence measured values by the cell vendor of Isc and Voc were 0.4762 A and 2.71 V respectively.Minimal degradation is indicated.The small reduction in Isc value observed is likely due to UV radiation degradation of the cover glass adhesive.The Voc results indicate also minimal degradation.The low degradations observed are expected due to the low predicted charged particle radiation environment of the spacecraft orbit and the low activity of present solar cycle.Solar array power analysis from flight data indicate that there have been no string losses to date.Telemetry resolution is not adequate to estimate individual cell short circuit failures.Fig. 9 and Fig. 10 show the minimum and maximum temperatures during an orbit of the +Y and -Y wing panels observed to date.The temperatures extremes are less than the thermal predictions.

Figure 8 .
Figure 8. Voc and Isc variations during the mission

Figure 9 .
Figure 9. Plus Y panels minimum and maximum temperatures to date

Figure 11 .
Figure 11.Battery End of Charge & Discharge voltages and eclipse durations to date

Figure 14 .
Figure 14.Battery Voltage v.s.Normalized SOC Rp is the chemical ionic polarization/diffusion resistance due to lithium diffusion in the electrodes.Rp is time varying that typically takes 10 to 15 minutes to settle.The total internal resistance R in =R dc + Rp(t)

Figure 15 .Figure 16 .
Figure 15.Flight battery voltage and model results Figure 17 Model battery capacity degradation predictions and Rin (three batteries) variations to dateThe model is also being used to predict effects of special operations/manoeuvres including effects of eclipses by the moon.Figure18is an example of battery voltage model results and orbital data of a recent lunar partial shadowing during three orbits and the corresponding model predictions.Lunar shadowing percentage is also shown.The system level information required for the model are: battery charge voltage & current limits, orbital beta angle, battery & solar array temperatures and day of year for sun-earth distance.Only the instantaneous orbital spacecraft power is needed for the model to calculate the battery voltage, current and SOC.The start of eclipse discharge is used in model calculation as the starting point with initial values of voltage and SOC.Battery on-orbit internal resistance characterization is also being performed during full sun periods by utilizing the PSE capability for setting the charge current to a negative value.After stabilizing the taper current to zero at a selected operating voltage, the batteries are discharged at constant rates starting from same SOC.The internal resistance is derived as delta V/delta I.The Rin characterization were conducted during recent two full sun periods about 3 month apart.The differences were within measurement uncertainties during the period.