Spacecraft Eclipse Power Budget Project

This project builds a spacecraft eclipse power budget for a small Earth-observation spacecraft. The goal is to produce an engineering power-budget deliverable that closes energy storage, solar-array recharge, depth of discharge, degradation margin, safe-mode survival, and validation evidence.

The project is simplified but realistic enough to expose the system trade-off. A spacecraft power subsystem is not sized from one peak load. It is sized from operating modes, orbit timing, eclipse duration, battery limits, array degradation, power-conversion efficiency, thermal constraints, fault modes, and mission operations.

Project Objective

Create a mode-based power budget that answers:

1. How much energy must the battery supply during eclipse?
2. What battery nameplate energy is required at beginning of life to satisfy end-of-life depth-of-discharge limits?
3. How much solar-array power is required during daylight to run the spacecraft and recharge the battery?
4. Does the selected battery and array preserve margin in nominal, downlink, imaging, safe-mode, and degraded cases?
5. What validation evidence is needed before launch?

The final deliverable should include a load table, eclipse energy calculation, battery sizing, solar-array sizing, margin table, safe-mode check, telemetry limits, and verification plan.

Baseline Mission Scenario

Assume a small spacecraft in low Earth orbit.

Parameter	Value
orbital period	96\ \text{min}
daylight duration	61\ \text{min}
eclipse duration	35\ \text{min}
regulated spacecraft bus	28\ \text{V}
maximum allowed end-of-life DOD	35\%
end-of-life battery capacity factor	0.80
battery discharge efficiency	0.92
battery charge efficiency	0.90
solar-array aging factor	0.82
pointing, temperature, and illumination factor	0.85
power-conversion efficiency from array to bus	0.92
design contingency on eclipse energy	20\%

These values are screening assumptions. A real spacecraft power budget must use orbit analysis, attitude profiles, cell vendor data, battery qualification limits, heater control logic, radiation environment, thermal-vacuum evidence, harness losses, converter efficiency maps, deployment risk, and mission operations rules.

Step 1: Build the Mode Load Table

Use a mode table rather than one average load.

Load	Daylight standby	Imaging	Downlink	Eclipse
onboard computer and avionics	8\ \text{W}	8\ \text{W}	8\ \text{W}	8\ \text{W}
attitude control	10\ \text{W}	18\ \text{W}	10\ \text{W}	10\ \text{W}
payload	5\ \text{W}	35\ \text{W}	5\ \text{W}	5\ \text{W}
communication system	4\ \text{W}	4\ \text{W}	29\ \text{W}	4\ \text{W}
thermal control	4\ \text{W}	4\ \text{W}	4\ \text{W}	12\ \text{W}
power and telemetry overhead	0\ \text{W}	6\ \text{W}	0\ \text{W}	0\ \text{W}
total	31\ \text{W}	75\ \text{W}	56\ \text{W}	39\ \text{W}

Assume each orbit includes:

• 61\ \text{min} of daylight baseline operation;
• 12\ \text{min} of imaging during daylight;
• 8\ \text{min} of downlink during daylight;
• 35\ \text{min} of eclipse.

Imaging and downlink are treated as load increments above daylight standby, not as separate full-daylight intervals.

Step 2: Calculate Eclipse Energy

Eclipse load is:

Eclipse duration is:

Bus energy required during eclipse is:

Apply the design contingency:

Engineering Comment

The contingency is not a substitute for thermal analysis. It covers early sizing uncertainty, not uncontrolled heater duty cycles, unmodeled converter losses, or a safe-mode attitude that increases eclipse heater load.

Step 3: Size Battery Nameplate Energy

The battery must deliver the contingency-adjusted eclipse energy at end of life while staying below the maximum allowed depth of discharge.

Usable bus energy from a battery with beginning-of-life rated energy E_{BOL} is:

Required beginning-of-life battery energy is:

Substitute:

Select:

End-of-life usable bus energy is:

Eclipse energy margin is:

Percentage margin relative to the required contingency energy is:

Engineering Comment

The selected battery is not oversized simply because nominal eclipse energy is 22.75\ \text{Wh}. The design must survive capacity fade, allowed DOD limits, discharge efficiency, load uncertainty, and operations constraints at end of life.

Step 4: Check Actual End-of-Life Depth of Discharge

For the selected battery, actual end-of-life DOD needed for the contingency eclipse energy is:

So:

This is below the limit:

Engineering Comment

This check is important because a nameplate battery capacity does not state mission usability. The usable energy depends on end-of-life capacity, allowed DOD, converter losses, temperature, current limit, and reserve policy.

Step 5: Check Battery Current and C-Rate

At the 28\ \text{V} bus, approximate end-of-life ampere-hour capacity is:

Eclipse current is:

Approximate C-rate is:

Peak imaging plus downlink power may reach:

Peak bus current is:

Peak C-rate screen:

Engineering Comment

The battery passes the first C-rate screen, but final acceptance still needs cell-level current limits, temperature limits, battery-management-system limits, fuse coordination, harness voltage drop, and transient load testing.

Step 6: Calculate Daylight Energy Demand

Daylight standby energy:

Imaging increment above standby:

Downlink increment above standby:

Total daylight bus energy is:

The battery recharge energy needed at the bus is:

Total energy the array must support during daylight is:

Average required bus power during daylight is:

Engineering Comment

The array must do two jobs during daylight: run the spacecraft and replace the energy removed during eclipse. A design that covers only instantaneous daylight loads can slowly walk the battery state of charge downward over repeated orbits.

Step 7: Size Solar-Array Beginning-of-Life Power

Effective end-of-life bus power from a beginning-of-life array power P_{array,BOL} is:

The combined factor is:

Required beginning-of-life array power without design margin is:

Apply 20\% sizing margin:

Select:

Expected end-of-life effective bus power is:

Power margin relative to the daylight recharge requirement is:

Engineering Comment

The selected array has energy margin, but it still needs deployment, thermal, radiation, contamination, shadowing, attitude, and power-converter validation. Solar-array nameplate power is not the same as available bus power during the actual orbit.

Step 8: Check Safe-Mode Eclipse Survival

Assume safe mode reduces nonessential loads but increases heater conservatism.

Safe-mode load	Power
onboard computer and fault management	7\ \text{W}
receiver and beacon	5\ \text{W}
attitude safe-mode sensors/actuators	5\ \text{W}
heaters	14\ \text{W}
total	31\ \text{W}

Safe-mode eclipse energy is:

With 20\% contingency:

This is below selected end-of-life usable energy:

Engineering Comment

Safe mode is not always lower power than nominal mode. It may turn off payloads, but it can also increase heater duty cycle, communications beacon use, or attitude-control power. A safe-mode budget should be calculated, not assumed.

Step 9: Define Telemetry and Operating Limits

The project should hand operations explicit limits:

Signal	Watch threshold	Action threshold
battery SOC before eclipse	55\%	50\%
predicted eclipse DOD	30\%	35\%
battery temperature	mission-specific	inhibit high load outside qualified band
array bus power in daylight	below 85\ \text{W} sustained	review attitude, deployment, contamination, or degradation
heater duty cycle in eclipse	above budget by 20\%	enter thermal power review
repeated negative orbit energy balance	any trend over several orbits	reduce payload duty cycle or downlink

These limits should be tied to flight rules. A warning that does not trigger an operating decision is not a complete power-management control.

Validation Evidence

The project deliverable is not complete until the following evidence exists:

1. mode load measurements with calibrated equipment;
2. battery capacity test at relevant temperature and discharge rate;
3. charge and discharge efficiency evidence;
4. solar-array current-voltage test and deployment verification;
5. power-converter efficiency map across expected loads;
6. thermal-vacuum test or correlated thermal model for eclipse heater duty;
7. radiation and aging assumptions for end-of-life capacity and array output;
8. software-in-the-loop or hardware-in-the-loop power-mode transition test;
9. safe-mode power and communication test;
10. operations procedure for SOC limits, load shedding, and payload duty-cycle reduction.

Final Decision

The screened design can proceed to detailed review with:

• 140\ \text{Wh} beginning-of-life battery energy;
• 150\ \text{W} beginning-of-life solar-array power;
• end-of-life eclipse DOD of about 26.5\% under contingency;
• daylight recharge margin of about 28\ \text{W} at the bus after aging and conversion factors;
• required validation of thermal, battery, array, converter, and operations assumptions.

The main engineering lesson is that spacecraft power budgets are closed over modes and orbits, not over isolated component ratings. A credible design states when energy is used, when it is restored, what degrades, what margin remains at end of life, and what evidence proves that the power balance will hold in flight.