energy balance leo satellite

2nd IEEE International Conference on Power and Energy (PECon 08), December 1-3, 2008, Johor Baharu, Malaysia

Energy Balance Analysis of Small Satellite in Low Earth Orbit (LEO)

Sung-Soo Jang* and Jaeho Choi**

* Chungbuk National University, Cheongju, S. Korea. Email: [email protected] ** Chungbuk National University, Cheongju, S. Korea. Email: [email protected]

Abstract In general, satellite electrical power system plays an important role in its mission performance. To perform the mission successfully, the satellite should be provided with the sufficient power until End-of-Life (EOL). To supply the sufficient power, it is not easy to install of larger power generation and storage source with power control units on satellite. Also, the thermal control is difficult in the space. The various technologies have been developed to reduce the satellite power sizing and to design power efficiently. The peak power tracking (PPT) method has been developed for spacecraft power system to use maximum available power of solar array. In general, to compensate for the solar cells degradation at EOL in satellite, the solar array will be designed with big enough margin. And, when the spacecraft exits eclipses, the peak output power of a solar array becomes very higher than its peak output power in the normal sunlight operation power. As a result, the spacecraft power system should dissipate high excess power through shunt regulators in the beginning of mission and after eclipse exit. To avoid the high power dissipation in the beginning of life and to utilize maximum available power after eclipse exit, the PPT is used in many recent spacecraft. Particularly, this PPT method is useful for the spacecraft in lower orbit because it goes through a large number of eclipse.

Keywords Electrical power system (EPS); Satellite; Peak power tracking (PPT); End-of-life (EOL)

I. INTRODUCTION The purpose of this paper is to present the results of

trade-off study to decide the sizing of Electrical Power System (EPS) in small satellite in Low Earth Orbit (LEO). Also this paper describes the power analysis methods and input data used in the EPS End-of-Life, worst case, and energy balance analyses for payload operations of satellite mission in LEO. Both 01:10 PM and 01:25 PM crossing time are considered, so the required power in each case is analyzed with satellite roll maneuver according to payload operation concept.

Cases considered during the analysis effort were each payload roll maneuvers degrees for Local Time Ascending Node (LTAN) of 01:10 PM and 01:25 PM. In addition, data transmission to the Ground Station during eclipse is investigated at LTAN of 01:00 PM, to assess the greatest science mode battery DOD.

II. ELECTRICAL POWER SYSTEM The EPS provides solar energy, electrical energy

conversion and storage, voltage conversion, power

control and distribution. The solar array (SA) will be capable of supplying adequate power to recharge the battery and maintain energy balance in each orbit at the end of life. The rechargeable battery will be capable of supplying steady state and transient electrical power to spacecraft loads and to the instrument throughout the various injection or operational modes. It provides control circuits and driving motors to give single axis pointing control for the solar array. It develops and conditions telemetry measurements for selected power system monitoring points. The EPS provides redundant control of all the deployment release circuits. It also provides the DC wiring harnesses for electrically interconnecting the spacecraft and instrument equipment.

The EPS functional block diagram is shown in Fig. 1. The SA generates the electrical power during periods of solar illumination throughout the operational life of the spacecraft, delivers the sufficient power to supply normal spacecraft bus, payloads the power demands, and achieves the per-orbit energy balance.

The battery will provide electrical energy to the satellite during pre-launch operations, the launch phase, eclipse periods, and during periods of peak power demand that exceeds solar array output capability.

The solar array regulator (SAR) converts power from the spacecraft solar array to the battery-clamped primary DC power bus.

The power control unit (PCU) receives primary power from the SAR during the sunlight portion of the orbit or the battery during eclipse and creates a primary power bus. The primary bus power is used for battery charging, and to provide power to the spacecraft loads. And the low voltage converter in PCU is supplied from the main power bus.

Fig. 1. EPS functional block diagram.

1-4244-2405-4/08/$20.00 2008 IEEE 967


III. The Required Power for Satellite Mission

A. Assumptions In this paper, 700km altitude, the sun-synchronous

orbit is used for satellite in LEO. Before anything else, to perform the power analysis, some assumptions are taken for all LTAN with roll maneuver to take a consideration of worst case conditions.

The assumptions are as followings: - EOL worst case solar array degradation - Summer solstice intensity - Battery has one shorted cell - Solar array high temperature - Spacecraft load profile with payload operation - Solar array offpoint angle variation - Solar array shadow is present during most of suntime,

shadowing solar cell strings will not producing output current

- Solar array clamped to battery voltage after eclipse

B. Considered Input Data for Power Analysis In this paper, only the science mode power analysis

was considered. In satellite power analysis, of course, all environmental conditions which can affect the SA power are considered. So, in the paper, all factor that affect the power including the load profile, SA temperature, SA offpoint angle, and SA shadowing effect will be discussed. Also, margin factors are included in the energy balance analysis program for worst case conditions.

1) Satellite Load Profile Fig. 2 shows the load profile including delta attitude

slew power during 20% Normal Mission and 10% with 30deg. roll maneuver, 2% with 56deg. roll maneuver, and 2% with 30deg. pitch maneuver in science mode. Added attitude slew power is needed to driving the reaction wheel in roll and pitch maneuver. Maximum satellite load power will be around 780 watts during payload operation.

2) Solar Array Temperature SA temperature is considered in this analysis. Of

course, the temperature margin uncertainty is added to thermal engineers temperature data to consider the worst case conditions. Fig. 3 presents the SA worst-case hot temperature profile with temperature margin uncertainty added during orbit suntime at EOL.

3) Solar Array Shadowing Effect The shadowing effect has been analyzed in the SA as

an orbit suntime function. In worst case analysis, supposed that even though the string is shadowed partially, this string cant produce the power. So, in 01:25 PM crossing time with roll maneuver, partially shadowing strings cant produce the power in the analysis. Fig. 4 is SA shadowing effect in 01:10 PM and 01:25 PM crossing time.

4) Solar Array Offpoint Angle In the roll maneuver operation, systems engineering

has considered the each roll angle case at each crossing time. In the zero roll operation at 01:25 PM crossing

time, a worst case of pitch-axis error and computed the value of equivalent yaw error as the difference between the real and nominal -angle occuring at summer solstice were calculated as SA offpoint angle. The difference between the actual -angle and the nominal -angle provides the second orthogonal component of the SA offpoint angle. But, in the each roll maneuver operation the offpoint angle used is provided by system engineering and SA drive error angle from attitude control engineer. Figure 5 is the SA offpoint angle as a function of orbit suntime.

C. PPT Operation Characteristic Peak Power Tracking (PPT) algorithm is used to

(a)

(b)

Fig. 2. Satellite load profile in science mode. (a) 20% mission load profile in science mode, (b) Mission load in 10% with 30deg. 2% with 56 deg. Roll, and 2% with 30deg. Pitch maneuver.

Fig. 3. Solar array temperature profile EOL worst-case science mode.

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(a)

(b)

Fig. 4. Solar array shadowing effect. (a) SA shadowing factors in 10% with 30deg. and 2% with 56deg. roll mission @01:10PM: (b) SA shadowing factors in 10% with 30deg. and 2% with 56deg. roll mission @01:25PM.

extract the maximum SA power for satellite-load and high-rate battery charging power. The use of PPT algorithm results in EPS optimal sizing to support the satellite mission successfully.

In Fig. 1, Solar Array Regulator (SAR) consists of buck regulator with pulse-width-modulation (PWM) duty cycles (D) by processor commands. With this architecture, PPT algorithm is operating to track the maximum solar array power. The calculation of PWM duty cycles for PPT is based on the transfer function of a buck-converter. General buck-converter output voltage (Vout) has linear relation to the PWM duty cycles (D) and the input voltage (Vin). The buck-converter is directly connected to the solar array, as shown in Fig. 1. The transfer function characteristic of a buck-converter is combined with the solar array operation characteristic curve.

The solar array has specific I-V curve characteristic with illumination, temperature, and degradation. As temperature increases, the solar-array open-voltage (Voc) decreased dramatically. But it shows that solar-array short-circuit current (Isc) is not changing much.

Fig. 6 is the ratio (Kmp) of open-circuit voltage (Voc) and peak-power voltage (Vmp) on orbit. The Kmp is almost constant during sunlight on orbit in Fig. 6. The SAR transfer function is driven as following Eq. (1):

SAR Duty Cycles for PPT = 255 X [Vbus / (Voc/Kmp) (1)

(a)

(b)

Fig. 5. Solar array offpoint angle. (a) SA offpoint angle in 10% with 30deg. and 2% with 56deg. roll mission @01:10PM: (b) SA offpoint angle in 10% with 30deg. and 2% with 56deg. roll mission @01:25PM. where, Kmp = Voc / Vmp.

The Kmp for the maximum solar array power varies from BOL to EOL. And calculated PWM duty cycles with Kmp for SAR operation will be used to track the solar array peak power region. To update Kmp in real operation, the solar array data should be analyzed and the new value should be updated. To calculate PWM duty cycles for SAR operation to track the peak power of solar array, there are several steps in PPT algorithm.

Step 1) Read the battery voltage (Vbat) Step 2) Command SAR1 PWM duty cycle to zero Step 3) Read SAR1 solar array open-circuit voltage

Fig. 6. Kmp characteristic curve in PPT algorithm.

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Step 4) Calculate the new PWM duty cycle (D) for PPT D = Kmp x Vbat / Voc

Step 5) Command the new PWM duty cycle to SAR

The PPT algorithm to find maximum solar array power using Eq. (1) is guaranteed in the range of normal charging battery voltage. In the lower voltage of the battery caused by some faults, the PPT algorithms can not find solar array peak power efficiently. But the long time duration of lower battery voltage could not be happen in satellite power design. If any, the time is a few milliseconds like as fuse blown in the PCU, freewheeling diode and MOSFET short in the SAR.

Figure 7 shows the maximum SA power voltage (Vmp) in PPT mode using EOL solar cell characteristic on orbit. The PPT constant value for EOL power simulation was applied to track maximum SA power. After eclipse, Vmp reached about 50 volts and it decreased until around 40 volts in PPT operation mode during sunlight. In the PPT operation mode, the SAR PWM duty cycle was analyzed using power simulation.

And, it was illustrated in Fig. 8. After eclipse exit, PWM is operating at 57.6% to track maximum power and is going up till 75.1% in higher SA temperature. Fig. 9 shows the normalized solar array output with respect to SAR switching duty cycles at 28deg.C of solar array.

D. Analysis Method The suntime-dependent input data sets previously described were entered into the EPS energy balance computer program. The output of this program includes an identification of the minimum amount of solar array power required to achieve per-orbit safe battery recharge, as well as the battery depth-of-discharge incurred during eclipse.

IV. CONCLUSION A positive SA power margin exists for all satellite

operation cases analyzed. Table 1 summarizes the EPS energy balance analysis results in each crossing time with roll maneuver. As shown in Table 1, the power margin is positive and battery Depth of Discharge (DOD) is acceptable in all cases. Minimum power margin of 4.9 % occurs at 01:25 PM crossing time with the largest roll maneuver. And battery DOD range of 26.35 % to 26.67% occurs for all roll maneuver cases. The maximum battery DOD is 33 % for Korean ground station pass during eclipse with LTAN of 01:00 PM. Fig. 10 shows that the satellite battery cycle capability with a 33 % DOD, providing 81 % cycle life margin.

Figure 11 shows that the results of required power to provide the satellite mission with LTAN. The required power of 01:10PM is larger than that of 01:25PM. Figure 12 presents SA performance and battery charging power of 10% mission with 30deg. Roll maneuver operation in science mode with largest roll and shadow at 01:25PM crossing time. Battery may be discharged

during the roll maneuver because Fig. 12 uses the minimum value of solar array power needed for achieving energy balance in one orbit. The actual SA predicted power capability in orbit at EOL is larger than the required power. If this power is produced, even though in the worst case orbit, the discharging current in the battery will not occur.

In PPT operation, the peak voltage of SA is around 47 volts. Due to the EPS operation mode, the SA input operation point will be adjusted to supply the proper power of the satellite load.

In the eclipse, battery voltage will be down below the 25 volts, but it will be reached max 32 volts in carging mode. And maximum charging current is approximately

Fig. 7. SA Vmp in PPT operation.

Fig. 8. SAR PWM characteristic curve in PPT operation.

Fig. 9. Normalized Solar Array Output vs. SAR Duty Cycle @28deg.C

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Table 1. EPS energy balance analysis results.

Crossing Time 01:10 PM

Satellite Operation

Normal Roll Pitch 20% 10% 2% 2% 0deg. 30deg. 56deg. 30deg.

SA Power Generation(W) 1,050 1,050 1,050 1,050

Required Power(W) 984 946 995 820

Power Margin(%) 6.7 10.9 5.5 28

DOD(%) 26.67 26.67 26.67 26.67 Crossing Time 01:25 PM

Satellite Operation

Normal Roll Pitch 20% 10% 2% 2% 0deg. 30deg. 56deg. 30deg.

SA Power Generation(W) 1,050 1,050 1,050 1,050

Required Power(W) 988 950 1001 837

Power Margin(%) 6.3 10.5 4.9 25.4

DOD(%) 26.35 26.35 26.35 26.35

Fig. 10. Battery life cycles prediction with DOD.

Fig. 11. Required power for science mode operation in 01:10PM and 01:25PM.

25 amps in PPT operation. Battery DOD in previous orbit was 21.9%, but battery will be fully recharged in the next sunlight and Recharge Ratio (RR) will be reached 1.12 before next eclipse entrance.

(a)

(b)

Fig. 12. Satellite load power, solar array, and battery performance prediction with minimum required power.

REFERENCES [1] W.J.Larson and J.R.Wertz, Space Mission Analysis

and Design, Kluwer Academic Publishers, 1995. [2] Willard R. Scott, Sealed Cell Nickel Cadmium Battery

Application Manual, NASA, 1979. [3] Mukund R. Patel, Spacecraft Power Systems, CRC

Press, 2005.

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energy balance leo satellite

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