3.6.21 · D5Spacecraft Structures & Systems Engineering
Question bank — Spacecraft bus — structure, power, thermal, ADCS, C&DH, comms, propulsion
True or false — justify
The spacecraft bus does the science, the payload just supports it.
False — it is the reverse. The payload is the mission (camera, telescope, sensor); the bus is the skeleton-plus-life-support that keeps the payload alive, pointed, and heard.
A CubeSat with 1.63 mm² of load-bearing wall area is safe because that satisfies the yield-stress calculation.
False — yield strength only checks that the metal does not permanently deform. Real walls are 1–2 mm because stiffness (vibration modes) and safety factors dominate, not raw strength.
If a structure survives the axial launch load it will survive the whole launch.
False — steady axial g-force is only one load. Vibration, acoustic, and shock loads excite resonant modes and often size the structure more than the static load does.
The structural efficiency tells you which material is strongest.
False — it tells you strength per unit density. A denser material can be stronger yet worse for spacecraft, because every kilogram costs launch mass.
In space a spacecraft loses heat by radiation, convection, and conduction to the surrounding gas.
False — there is no surrounding gas, so convection to space is zero. Radiation () is the only path off the vehicle; conduction only moves heat within it.
A black surface (high ) is always a bad choice for a satellite.
False — high absorptivity is only bad when solar input is the problem. In a cold, sun-starved orbit you may want to absorb sunlight, so the "right" coating depends on the thermal environment.
Increasing emissivity always cools the spacecraft down.
False — appears in both the radiated output and the Earth-IR absorbed input (). Raising it boosts rejection but also intake, so the net effect depends on which term dominates.
Solar array power at end of mission equals the power at launch.
False — cells degrade (~2.5%/yr in LEO from radiation). You must size for Beginning of Life so that after degradation there is still enough at End of Life.
If the battery round-trip efficiency were 100%, the solar array would only need to supply the average load.
False — even with perfect batteries the array must generate the whole orbit's energy during only the sunlit fraction, so it still exceeds by the eclipse/sun time ratio.
Spot the error
" radiates the heat away."
The temperature must be to the fourth power: . Radiated power scales as , which is why a small temperature rise dumps a lot more heat.
" in the Stefan–Boltzmann equation is in degrees Celsius."
Wrong — it must be absolute temperature in Kelvin. Using 20 instead of 293 K underestimates the radiated power by a factor of tens of thousands.
"Solar input equals where is the whole surface area."
Only the projected area facing the Sun absorbs sunlight, not the total surface. A cube presents one face to the Sun, not six.
"Eclipse fraction 35% of a 95-min orbit means eclipse lasts 35 min."
min, so about 33 min, not 35. The 35 is a percentage, not a duration.
"Since , use the largest area you can to be safe."
The inequality gives a minimum; oversizing wastes mass. The engineering goal is the smallest area meeting strength and stiffness, because mass is the enemy.
"Albedo heating comes from the Sun, so it disappears during eclipse."
Albedo is sunlight reflected off Earth, so on the day side it exists — but during eclipse the spacecraft is behind Earth's shadow, so albedo does vanish there. The trap is calling it a Sun-independent term; it tracks illumination of Earth's surface below.
Why questions
Why does the structure typically get sized by the launch environment, not the on-orbit environment?
Because launch delivers the harshest mechanical loads — high g-forces, vibration, and acoustic shock — while orbit is nearly force-free. Survive launch and orbit is easy.
Why can't we just make the solar array exactly the size of the average load?
The array only produces power in sunlight, but the load runs all orbit. It must overproduce during the sunlit part to bank energy for eclipse, plus cover battery losses.
Why do we care about the ratio rather than each value alone?
The equilibrium temperature depends on the balance of solar-in (via ) versus radiated-out (via ). A ==low == stays cool in sunlight; a high one stays warm — so the ratio sets the running temperature.
Why does adding a safety factor multiply the structural requirement by 10–20×, not just 2×?
Because several independent unknowns stack up: material scatter, vibration/resonance, handling, thermal cycling, and modelling uncertainty each demand margin, and they multiply rather than add.
Why is the battery needed at all if the solar array can meet the load?
The array only works in sunlight; during eclipse there is zero generation, so the battery must carry the entire load. Storage bridges the dark half of every orbit.
Why is emissivity split into two separate roles in the thermal balance?
The same surface both radiates the spacecraft's own heat away () and absorbs Earth's infrared glow (). One coating property drives both directions of IR exchange.
Edge cases
What happens to the power budget for a spacecraft with zero eclipse (e.g. a dawn–dusk sun-synchronous orbit)?
, so the sizing term vanishes and — the array only needs to cover the load, and the battery is nearly idle.
What does the thermal balance predict for a spacecraft in deep space, far from Sun and Earth?
The solar, albedo, and Earth-IR inputs all drop toward zero, leaving only internal dissipation to radiate. Equilibrium temperature falls, so the design problem flips from cooling to keeping warm (heaters).
What if internal dissipation were the only heat source (fully shaded orbit)?
Then alone, and the equilibrium temperature is set purely by waste heat and radiating area — no term appears at all.
What happens to required array power as eclipse duration approaches a full orbit ()?
The ratio , so — an unphysical demand meaning no realistic array can survive a near-permanently-eclipsed orbit; you'd need a different power source.
For a surface with (grey body), how does its equilibrium temperature in sunlight compare to Earth's coating-driven tuning?
With the ratio is 1 and the coating gives no thermal tuning leverage; the temperature is fixed by geometry and flux alone, which is why engineers deliberately pick coatings where .
What is the minimum wall area if launch acceleration were zero (imaginary gentle launch)?
The static formula gives , i.e. strength imposes no constraint — but the structure still needs finite thickness for stiffness, handling, and thermal loads, showing the yield calc alone never sets a real minimum.
Recall Quick self-test
The one property that is zero in space and therefore never appears in spacecraft thermal balance ::: Convection (no surrounding gas to carry heat away). The environment that usually sizes the primary structure ::: Launch — g-forces, vibration, acoustic, and shock loads. Why the solar array is always bigger than the average load ::: It must bank eclipse energy and cover battery losses while only generating in sunlight.