Estimate orbit speed and period, sunlight, solar array area, battery capacity and radiator requirements with transparent educational models.
Change altitude and computing load to explore the power and cooling behind computing in orbit. Review the assumptions, equations and sources before using the estimates. Kaveri Orbital is a personal passion project by Arun Govindappa and his son, Naren Arun, exploring orbital mechanics, space technology and computing beyond Earth. Naren is completing his degree in astrophysics at Pennsylvania State University (Penn State).
Penn State: Department of Astronomy and Astrophysics
Calculated with Kaveri Orbital model 0.1.0, not a NASA design: 100 kW IT plus 20 kW support load; beta 0°. The circular orbit takes 95.5 minutes at 7.59 km/s, with 60.0 minutes of sunlight and 35.6 minutes of eclipse. Estimated solar array footprint is 899 m², nominal battery capacity 140.3 kWh and radiator emitting area 311 m². These are illustrative model outputs, not hardware specifications.
Solar flux 1,361 W/m²; cell efficiency 30%; cell packing 90%; pointing factor 95%; array end-of-life output 80%; bus conversion efficiency 95%. Battery charging and discharging efficiencies each 95%; usable depth 80%; retained capacity 80%. Sizing margin 20%; radiator surface 50°C; IR emissivity 90%; absorbed environmental heat 50 W/m². Advanced assumptions in the Explorer can be changed. The example remains fixed so it can be reproduced.
In a circular Earth orbit, altitude determines speed and period in this two-body model. Higher circular orbits take longer and have a lower orbital speed. Here r is Earth radius plus altitude in kilometres, μ = 398,600.4418 km³/s², v is km/s and T is seconds. Earth radius is approximated as 6,371 km.
v = √(μ / r); T = 2π √(r³ / μ)No. Eclipse duration depends on orbit geometry and the solar beta angle, not altitude alone. Beta is the angle between sunlight and the orbital plane. In this cylindrical-shadow model, some high-beta geometries have no eclipse; beta changes with orbit orientation and season. Penumbra and orbital evolution are omitted.
Start with continuous IT and support power, then include the energy needed to recharge the eclipse battery during sunlight. Conversion losses, cell efficiency, packing, pointing, end-of-life degradation and a sizing margin increase the required array area. This Explorer reports module footprint after packing; it does not size deployment structures. In the equation, Parray is required sunlit array power in watts, S is 1,361 W/m² and each efficiency or reduction factor is a fraction.
Aarray = Parray / (S × cell efficiency × packing × pointing × retention)NASA: Small Spacecraft Power · Kaveri Orbital: equations, assumptions and limits
Power multiplied by eclipse hours gives delivered energy in kWh. Nominal battery capacity must also account for discharge losses, usable depth of discharge, retained end-of-life capacity and a sizing margin. Charging losses are included separately when sizing the solar array. Battery mass, cycle life and thermal regulation require further analysis. Pload is continuous bus load in kW; margin is 1 + the entered percentage divided by 100, and efficiencies are fractions.
Battery kWh = Pload × eclipse hours × margin / (discharge efficiency × usable depth × retained capacity)NASA: Small Spacecraft Power · Kaveri Orbital: equations, assumptions and limits
A vacuum provides no ambient air for convective cooling. Heat must reach radiating surfaces through the spacecraft thermal system and then leave as thermal radiation. Radiator temperature, emissivity and absorbed sunlight, Earth infrared and albedo affect net rejection. A cold background does not remove the need for radiators and heat transport.
The model divides average onboard heat, with sizing margin, by net emitted heat per square metre. T is radiator surface temperature in kelvin; ε is infrared emissivity; σ = 5.670374419 × 10⁻⁸ W/m²/K⁴; qabsorbed is the assumed absorbed environmental flux in W/m². Qheat is in watts and margin is 1 + the entered percentage divided by 100. A non-positive net flux is marked infeasible. Emitting surface area is different from panel footprint, and this model does not solve shading, view factors or transient peaks.
A = Qheat × margin / (εσT⁴ − qabsorbed)NASA: Small Spacecraft Thermal Control · Kaveri Orbital: equations, assumptions and limits
No. These educational estimates do not include a complete spacecraft, launch costs, radiation reliability, communications capacity, maintenance, redundancy or a commercial operating model. Model checks are not independent engineering certification. Verify consequential decisions with qualified specialists.
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