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From launch to end of a mission, the Thermal Control Subsystem engineers task is to retain the temperature of all spacecraft components, subsystems and the total flight system within specified limits for all flight models. As a result of there being no atmosphere or gravitational field, convection does not subsist for unmanned spacecraft but, conduction and radiation are the heat transfer mechanisms affecting spacecraft thermal control. Gradient temperatures and specific stability limits are to be imposed on flight system elements in some cases. In general the thermal control subsystem's mass power and control requirements are below 10% of the total flight system resources
Active techniques require the maneuver of spacecraft resources, including electrical power, sensing and data storage, data handling and control. Active thermal control methods use heating or cooling thermal transfer, variable rejection and sensing data. Active thermal control hardware includes coolers (sterling/sorption), thermal transfer (heat pipes/pump loops), thermal louvers, thermal switches, controllers, thermostats, heaters (electric/radioisotope), sensors (thermistors/PRT's (Platinum resistors thermometers)/thermocouples), dewars. Therefore, Active thermal control techniques require additional testing and consume a greater risk and cost. In the case of passive techniques being unable to deliver the control needed, Active techniques are used.
An initial mass budget allocation was set to thermo sub systems to be less than 25 kg's, and a maximum power requirement of 18 W. As a result of this low mass and power budget allocation it is increasingly likely that we use Passive Thermal Control technique in order to be within these constraints. Also, the simplicity, reliability and lower cost of this technique results this method undoubtedly being a more precise and appropriate choice for the lunar mission. The table shown below represents the passive thermal control hardware options allocated for the lunar mission for both chemical and electrical spacecrafts.
As you see above the loop consist multiple evaporators where the evaporator has its own compensation chamber. Each evaporator has a primary wick of an outer diameter 6.35 mm. there are two condensers attached to the two suggested radiators and also two thermo electric converters attached to the compensation chambers that are connected to the evaporator via flexible thermal straps. The flow regulator is located downstream of the condenser and also the coupling block is connected via the vapor line and the liquid line to the flow regulator. For the working fluid it's been decided that ammonia fluid will be used.
The loop heat pipes, utilizes boiling and condensation of the fluid to transfer heat, surface tension forces developed from the evaporator wick to circulate fluid 1-2. This passive process is self-regulated, in that liquid is drawn by the evaporator as necessary to be converted to vapor according to the heat applied. The two evaporators are placed in parallel in where both will work passively and no control valves are needed to distribute the fluid along the evaporators. These two evaporators will produce vapor which has the same temperature regardless of their own heat loads. The spacecraft internal instruments will be placed at their own optimum locations as the loop works as a thermal bus that provides a single interface temperature. None operational instruments will draw heat from the operational instruments as the evaporators will automatically share the heat among each other (2,6). Electrical heaters will furthermore not be needed for the spacecraft, as this loop will maintain instruments at their operating temperature. Instruments are able to operate independently without an affect from other instruments or affecting other instruments as of this passive and automated heat load sharing function. If at a point all instruments are turned off the entire loop could be shut downed keeping the compensation chamber above the minimum temperature.
Allowable instrumental temperaturecondenser/radiators will not receive any heat flow thus the loop work as a thermal switch. Keeping the primary wick at a outer diameter 6.35mm evaporator has reduce its mass more than 70% compared to normal where the overall heat pipes mass is just 0.8 kg .the small evaporators wick also reduce the required fluid inventory in the loop heat pipes.The overall mass and the volume of the thermal system has reduce. In a normal starter heater evaporator the power requirement is 20w to 40w but in this multiple evaporator system the startup power requirement is less than 5w the normal control heater on compensation chamber for temperature control its only cold buyers, heating only but has no cooling and the heater power is about 10w to 20w with the thermo electric convertor on the compensation chamber and the coupling block on transport lines for temperature control allows heating and cooling with a heating power variation from 1w to 5w.
Paragon space Development Corporation developed a new radiator teaming up with NASA Johnson space Centre developing a variable emissivity radiator, which was agreed for installment in this spacecraft. This radiator uses emissivity electrochromic technology unlike any other coating on the radiator, which has a constant emittanceelectrochromic films allow for variation of the radiator surface emissivity. This allows heat rejection variation in the space radiator that therefor is advantageous when compared to common radiators that vary the temperature to control the heat rejected.
Variable emissivity is a particular significance due to the power variation of the spacecraft during the lunar orbit arrival. We could also use a traditional spacecraft radiator but it has a quite a significant temperature swing which will be at a lower cost that is non-toxic and that uses water as its working fluid which is liable to freeze in lower power states. But emissivity radiator enables the heart to be rejected with a slight temperature swing .due to this highly sophisticated radiator it could eliminate the need of hazardous working fluids, two loop systems, heat pumps, heavy valves and some other heavy thermal associated components. This radiator is designed to be implemented on the exterior surface so any heat dissipation of the radiator itself will not interfere with the spacecraft interior. This is basically invented for active thermal controlling but due to the advantages and high efficiency this is used in the passive controlling system to be on the safer side to the lunar mission which will give thermal protection in a much higher state than required. Due to this implementation a huge mass is saved on the thermal subsystem.