GB2413861A

GB2413861A - Controlling power to satellite electric propulsion system during eclipses.

Info

Publication number: GB2413861A
Application number: GB0410056A
Authority: GB
Inventors: David John Milligan
Original assignee: VEGA GROUP PLC
Current assignee: VEGA GROUP PLC
Priority date: 2004-05-06
Filing date: 2004-05-06
Publication date: 2005-11-09
Also published as: GB0410056D0

Abstract

A control loop / algorithm is implemented on an electrically propelled spacecraft to estimate the remaining capacity in an energy storage device at a given time and automatically change the power utilisation of the electric propulsion system based on the estimate. Typically, this would be used to allow safe and efficient EP thrusting through eclipses on battery power, estimating and controlling the capacity of the batteries at the end of the eclipse. Also described is a power subsystem design trade off to allow thrusting through a subset of expected eclipses in orbit raising missions. The increased utilisation of the electric propulsion system allows for faster, and more efficient orbit raising. The faster orbit raising reduces radiation damage to spacecraft performing orbit changes through radiation belts. The gains produced could be used to increase the payload, reduce the launch mass, extend the lifetime, or any combination of the three.

Description

1 241 3861 Enhanced use of Spacecraft Electric Propulsion through eclipses

General Principle EP (Electric Propulsion) missions around a planetary body experience regular eclipses, depending upon season and the exact orbit. If solar energy is used, the energy storage system (usually batteries) will be sized to ensure safe spacecraft operation through the longest predicted eclipse. Since the EP load is typically a large fraction of the overall power budget, it may be decided to exclude EP usage in eclipses due to the mass penalty incurred by a large energy storage system. Also, thrusting through eclipses can be considered dangerous. Since EP systems are typically high power users, there is some danger that batteries will be rapidly and strongly depleted. This is an important consideration if the relative size of the battery capacity to EP power demand is low, especially when some failures are also taken into account.

The subject of this patent allows for safe and efficient thrusting through eclipses and increases the range of missions types where thrusting through eclipses can be performed. By increasing the confidence, acceptability and efficiency of EP thrusting through eclipses, orbit- raising missions can be completed quicker, using less propellant. In the case of performing orbit changes through radiation belts, this has the added benefit of reducing radiation exposure to spacecraft electronics, and reducing solar array degradation.

Part one of this patent concerns the differential sizing of total battery capacity and maximum discharge rate (e.g. size the maximum battery discharge rate to take EP loads, but do not size the total battery capacity to take the EP load through the maximum predicted eclipse length) . Also consider this when trading unregulated and regulated power buses for use on EP spacecraft. By designing in this way, the option of using batteries for EP is not removed. Even if battery sizing for the longest eclipses at full power is considered not possible, a subset of smaller eclipses could still be thrusted through. When this is considered in the mission analysis phase, propellant, and time, savings can be achieved. Also to be considered here is the effect of the necessary operational margins. If the spacecraft is designed such that it is unable to support any EP operation through eclipses then operationally it must be ensured that the EP system is off at eclipse entry. When this constraint is applied in conjunction with thrust level uncertainties, and relatively infrequent spacecraft commanding, the actual time the EP system is off around an eclipse can be several times the eclipse length (e.g. a 20 minute eclipse length can easily require a 100 minute EP off time).

Part two of this patent concerns the automatic control of EP power consumption, through eclipses, to target and control the battery capacity at the eclipse end. This part of the patent would be implemented using a software algorithm, although a hardware implementation of the logic is not excluded. The software algorithm computes the total power consumption at any one time. The software algorithm also estimates the total battery capacity at any time. The battery capacity can be estimated in any way.

Examples include, but are not limited to, the use of battery temperatures, voltages and currents for Lithium-Ion batteries, or the simple integration with time of battery currents and voltages, or the integration with time of the power consumption (from eclipse start) as computed in the first part of the algorithm above (coupled with an estimate of the charge level at eclipse start). The software combines the current battery capacity estimate with the current power consumption (including EP load), to estimate the battery capacity at the end of the eclipse. The eclipse length is also known by the software either because the software has an orbit model, or because the ground uplinks this data periodically. Finally, the end of eclipse battery capacity estimate is used to change and control the EP power consumption, such that a safe battery level is ensured at eclipse exit. Since the level is automatically targeted by OBSW, it is robust against failures, and at the same time, more efficient than could be done by ground control. If no OBSW control is used then this could only be implemented by ground prediction, which has to include much larger margins to cover failures that in most cases will not occur.

The EP system can be of any type by which propulsive force is generated by electrical power (e.g. including, but not limited to, stationary plasma thrusters, Hall effect thrusters, "ridded ion thrusters, Radio frequency Ion thrusters, magneto- plasmadynamic thrusters, arcjects, resistojets etc). The change of operating condition of the electric thruster(s) may be executed by changing any voltage and/or current, or any other characteristic that results in a change of specific impulse and/or thrust. An example would be to change the discharge voltage in a "ridded ion thruster to change the thrust, or the delta voltage on the extraction electrodes to change the thrust and specific impulse. Another example would be a change of discharge voltage and/or current in a stationary plasma thruster. In thrusters using applied magnetic fields this could also be varied to change the operating condition.

The energy storage equipment can be of any type. Batteries are often used in the descriptive text of this patent to ease understanding and flow, but the patent is applicable to any energy storage device. Examples include, but are not limited to, batteries, fuel cells and kinetic energy storage. The power subsystem architecture can be of any type, including, but not limited to, a fully regulated power bus, unregulated power bus, hybrid regulated/unregulated bus, maximum power point trackers etc. Although the next sections show specific examples of how the patent can be implemented they are merely examples and are not meant to be exhaustive. The patent covers all aspects of differential sizing of energy storage device capacity and maximum discharge rate to allow EP thrusting in some eclipses. The patent covers all aspects of any control loop / algorithm that changes the EP power consumption to target an end of eclipse battery (or energy storage device) capacity. The control loop / algorithm could also be used outside of eclipse if this were so desired (i.e. if the EP system is using battery power outside eclipse and a certain battery capacity is required to be guaranteed a certain time later). The control loop / algorithm could be implemented exclusively in hardware, exclusively in software, or by a combination of hardware and software.

Figure 1 shows an example of how the end of eclipse battery capacity EP throttling algorithm could be implemented. : 3

Example of differential sizing of maximum battery discharge rate and battery capacity to allow EP thrusting through some eclipses Early mission studies will identify a maximum eclipse length, which is given as an input to the power subsystem design. It is also possible that the natural maximum eclipse length as generated from the optimum trajectory is considered too long for the power system design (excessive mass of energy storage system) and this can then be used as a justification to alter the mission trajectory. In this example it is assumed a maximum eclipse length is given.

An example power budget is divided as follows: 1. Maximum eclipse length 2 hours 2. Nominal platform power consumption in eclipse (excluding EP) = 200W 3. Battery energy needed for longest eclipse = 400Whrs 4. To include margin, and cell failure, the required battery capacity = 400Whrs/0.6 = 667Whrs.

5. Expected battery discharge rate = 200W. To include margin allow 200W/0. 6 = 333W.

So in this example the batteries are sized only to support the nominal platform consumption, with some margin, through the worst-case eclipse (i. e. no EP thrusting envisioned in any eclipses). At the beginning of the mission the eclipses are actually only 20 minutes (e.g. an Earth centred GTO initial orbit with eclipses at perigee, that later incurs longer eclipses in an evolved orbit). In this case, only 67Whrs is required in the early eclipses for spacecraft nominal loads. If we target an 80% depth of discharge at the end of these early eclipses, then this leaves 467 Whrs for EP thrusting through the eclipse, or an average EP power consumption of 1400W.

This thrusting through early eclipses is only possible if the maximum discharge rate of the batteries is not sized in the same way as the battery capacity (i.e. do not limit battery discharge rate as in point 5 above). When including the necessary operational switch off margins, of around 40 minutes either side of the eclipse, this would enable an extra 100 minutes thrusting per orbit. In this example (orbit raising from GTO) this would initially allow an increase in allowable impulse per orbit of the order of 20%.

Example of an EP power control software algorithm for use through eclipses An example implementation software algorithm is shown schematically in figure 1.

The loop is activated if the spacecraft is in eclipse and the throttling algorithm is enabled (checks in stepl). In the next step, a computation is executed to predict the battery capacity at the end of the eclipse, based on an estimate of the present capacity, the current power consumption (including EP) and the eclipse length.

Two end of eclipse battery capacity target levels are known to the software (changeable by telecommand). 'Target_Low' is the first capacity level. If the EOE_BAT (end of eclipse battery capacity estimate) falls below this level (e.g. in steady state, an extra heater has switched on) the EP should be throttled down so that the end of eclipse battery capacity remains in the target range. The second level is Target_High'. If the EOE_BAT rises above this level (e.g. a heater is switched off), the EP power consumption can be safely increased. These two levels are placed sufficiently far apart such that rapid and repeated EP throttling is avoided (i.e. the difference between the two should be greater than the minimum power throttling step size capability of the EP system). These are the two checks performed in steps 3 and 9.

If the answer to step 3 is 'Y' then the EP system needs to be throttled down. It could also be possible that the EP system cannot be throttled down sufficiently to compensate, since EP systems typically have a minimum possible power consumption (EP_MIN). A check is performed in step 5 for this case. If true, the EP system is powered off (since even at minimum EP power consumption, the batteries will be too depleted at the end of the eclipse). If the answer to step 5 is 'N' then the EP can be throttled down, such that the end of eclipse estimate is again in an acceptable range (step 7). A new EP power level was calculated in step 4 (New EP_Pow), which produces an in range EOE_BAT called (Target_Mid) which is between Target_High and Target_Low. It will take some time for the EP system to achieve the new level, which is accounted for in the wait time of step 8.

If the answer to step 3 is 'N' and the answer to step 9 is 'Y' then the EP can be safely throttled up. The new EP level is calculated in step 10 and commanded in step 11. A wait time for throttling is included in step 12. If the answer to step 3 and step 9 is 'N' then no throttling is required, since the current spacecraft power consumption will produce an in range EOE_BAT. Step 14 sets out conditions to repeat the loop (i.e. still in eclipse, EP throttling still enabled and EP has not been commanded off).

Possible values for the variables mentioned above are shown below in table 1. The end of eclipse target levels could include some margin, such that the calculation made in step 2 can support 'one off' events (e.g. RW offloading - expected end of eclipse heater switch ONs). All values, and the exact nature of the flowchart, are indicative of a possible implementation of this patent, but they do not cover all possible implementations of EP throttling to control end of eclipse battery capacity.

Variabie Possible Value Meaning | Eclipse Flag True/False Spacecraft in eclipsed | EP_Throttle True/False EP throttling software loop based on EOE_BAT enabled? EP_ON True/False EP thrusting? EOE_BAT Estimate of the remaining battery capacity at the eclipse end Target_High 22% of full capacity Upper range of the desired end of eclipse battery capacity Target Low 18% of full capacity Lower range of the desired end of eclipse battery capacity New EP Pow. New power level for EP to be commanded.

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Target _Mid 20% of full capacity Target EOE_BAT that is used to compute New EP Pow

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EP_Min Minimum commendable power level for the EP system Table 1. Explanations and possible values for variables used in theflowchart in figure I

Claims

1. An electric propulsion control loop / algorithm, used to target an energy storage device capacity level after a given time, such that electric propulsion power utilisation is changed.

2. An electric propulsion control loop / algorithm, as claimed in claim 1, where the energy capacity at the end of an eclipse is targeted.

3. An electric propulsion spacecraft that uses differentially sized energy storage device capacity and energy storage device maximum discharge rate to allow the EP system to operate on the energy storage device through a subset of eclipses.

4. An electric propulsion control loop / algorithm, as claimed in the above claims, that is used to affect any operating characteristic of any electric propulsion unit.

5. An electric propulsion control loop / algorithm, as claimed in the above claims, that is implemented in software or hardware or by a combination of the two.

6. An electric propulsion control loop / algorithm and/or design, as claimed in the above claims, that is used to increase the efficiency and/or speed of any orbit change manoeuvres.

7. An electric propulsion control loop / algorithm and/or design, as claimed in the above claims, that is used to increase the efficiency and/or speed of any station keeping or attitude change manoeuvres.

8. An electric propulsion conko1 loop / algorithm substantially as herein described above and illustrated in the accompanying drawings.

9. An electric propulsion spacecraft using differentially sized energy storage device capacity and energy storage device maximum discharge rate as herein described above.