engineering

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The Thermal Control Techniques Engineering Essay

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.

Passive techniques are normally preferred as a result of their simplicity, reliability and lower risk and cost. These techniques are elements that require no further spacecraft resources as control, electrical power, and does not impose any data handling requirements once installed, it refrains from changing any property (such as heat rejection capability), it does not posses any external moving parts.

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.

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080032843_2008031434.pdf

Multilayer Insulation

(MLI)

Thermal Surfaces

Paints / Films / Chemical Coatings

Thermal Conduction Control

Isolation

Low Conductance Material

High Conductance

Al / Cu / Be / K1100 / APC

Thermal Storage

Mass of Materials

Phase Change Materials

Thermal Radiators

Normal Temperature

Cryogenic

Passive Cryogenic

Thermal Transfer

Heat Pipes

Radiation Block

Sun Shade / Planet Shade

Entry System

Heat Shield / Back Cover

Typical spacecraft design temperatures for both chemical and electrical spacecrafts.

Component /system

Operating system(Kelvin)

Survival temperature(Kelvin)

Digital electronics

273 k - 323 k

253k- 343k

Analog electronics

273k -313k

253k - 343k

Batteries

283k - 293k

273 - 308k

IR detectors

4k - 100k

4k - 308k

Solid state particle detectors

238k - 273k

238k - 308k

Momentum wheel

273k– 323k

253k – 343k

Solar panels

173k – 398k

173k – 398k

Above table represents spacecraft design temperatures for both systems’ components thermal limitation range in Kelvin. When designing both chemical and electrical space crafts its important that the space crafts temperatures lie within the survival temperature.as u can see from the above table, the operating temperatures exceed the survival temperatures. Therefore thermal-controlling systems should be allocated for almost most of the components of the two spacecrafts.

Heat pipes

Heat pipes are very efficient heat transfer devices on a continuous condensation of a working fluid cycle, inside a properly sealed tube. The processing power and electronic speeds, have given rise to unwanted heat within the spacecraft that needs dissipation by radiators into deep space.

Heat pipes will be attached to all the dissipating equipment’s to transport that heat to the radiators where the radiators will spread this heat within the radiators to process the mechanism of maintaining a stable thermal environment within the spacecraft. Smaller heat pipes will be integrated into electronic units for the cooling process of that unit.

Loop heat pipe system will be integrated into the spacecraft where heat from heat sources will be transported to the heat sinks while maintaining temperature control for all instruments under varying heat loads and environmental thermal conditions

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.

The fluid flow distribution among the 2 parallel condensers if passive and self regulating “2,6”. Conservation laws of energy mass and momentum are satisfied in the condenser section with the appropriate mass flow received to the condenser. When the radiator is exposed to a high warm environment the capillary regulator located down stream will prevent vapor entering the liquid return lines. Diversion of excess vapor flow will be done to the other condenser thus instruments will not be transmitted any heat from the hot radiator effecting the thermal diode action

http://mechanical-engineering.esa.int/thermal/heatpipes.html

http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080032843_2008031434.pdf

Radiators

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.

http://www.paragonsdc.com/brochures/VariableEmissivity_lo.pdf

The variable emissivity radiator systems consist of two radiators which wraps around thin film electrochromics and the system consists solid state control electronics for higher reliability controlling which in comparison with the old radiators this is the state of the art design.

In detail a heat rejection ratio greater than 8 – 1 is achievable and more could be achieved with allowance of temperature variations. Micrometeoroids orbital debris damages to the radiator could be avoided as the thin film elcrochromics acts as a protection layer for the radiator system. Electrochromic thin films developed by ‘Eclipse energy systems’ are less than 10 grams per square meter and this replaces the need for paints for the radiator system.emmitance state controller allows the radiator system to fix radiator emissivity in an intermediate state for better thermal controlling. The emmitance changes will be done by the state controller as programed for the voyage and the entire life span of the mission. A control voltage of ± 2V will be required for the system. This low power requirement is also a reason to choose this radiator for both chemical and electric space crafts as the radiators will be used on the loop heat pipes system which will give a much higher thermal protection for both systems which will protect the components with adequate protection. These radiator constructions are highly resistant to atomic oxygen equally to those old radiators.

http://www.paragonsdc.com/brochures/VariableEmissivity_lo.pdf

Above graph shows the variation in the thin film electrochromic’s thermal spectral emmitance between the low emissivity and high emissivity.

RADIATOR

Surface Material

OSR

Absorbance

0.6

Surface Area

1.28

m2

Emittance

0.72

Density

1136.4

Kg/m3

Mass of OSR

1.75

kg

Volume of OSR

0.00154

m3

Thickness

0.0012

m

Surface Material

AL surface withthin film electrochromics

Absorbance

0.92

Surface Area

1.28

m2

Emittance

0.64

Density

2700

Kg/m3

Mass

1.68

kg

Volume

0.000622

m3

Thickness

0.0004859

m

Surface material

Louvers cap

Absorbance

Variable

Surface Area

1.28

m2

Emittance

Variable

Density

2700

Kg/m3

Mass

0.89

kg

Volume

0.0003296

m3

Thickness

0.0002575

m

Total radiator mass

4.32

Kg

Radiator power

2

W

Multi layer insulation (MLI)

Multi layer insulation minimizes the radiative heat transferred from or to a spacecraft’s components transferred. Aluminized kapton or milar with thin net materials in between layering alternatively gives rise to this multilayer insulation. This layer is not conductive but radiative due to no primary air coupling between the layers. Larger multi layer insulation systems have low effective emissivity in general. Until all air has escaped from the MLI blankets, which take several hours experiments, could not be made in thermal vacuum tests. Fewer seams and penetration for unit area has increased the effectivity of larger MLI systems. During test and on orbit it precisely takes a long time for MLI systems air to be taken out. It’s decided to use MLI blankets in the spacecraft entire structure for thermal controlling other than the radiator panels and the engine nozzle. Even though the thermal system is generalized for both chemical and electrical spacecrafts need more MLI blankets to limit the heat input of the fuel tanks and the fuel lines as this consumes more area and volume. A surface area of 6.8 m2 of MLI blankets will be needed with a volume 0.0041 m3, thickness 0.0006 m ,  of 0.001, density 1348kg/m3 and a mass of 5.5 kg where in the electric spacecraft emissivity (),thickness, density is equal as to the chemical spacecraft but the surface area has decreased to 5.5 m2 , volume 0.0033m3 and the mass to 4.4484 kg.

Adhesive tapes

Adhesive tapes have been used as fillers to increase the heat transfer between the surfaces. These are made with thin sheets of metal oxides in elastomeric binders and these fill voids between contacting surfaces. Adhesive tapes are also used when manufacturing MLI. This is also used to modify the thermo optical property of a material or for a structural member by applying directly on them. Fillers create a thermal path critically but this property is difficult to test n the atmosphere.

Thermal Isolators

Thermal isolators are used to isolate instruments and other components from the spacecrafts body for low conductivity materials.

Thermo electric coolers

These cooler’s electrical current induces cooling of junctions between metals, which are dissimilar. These have relatively low efficiency.

Thermal Surfaces - Paints / Films / Chemical Coatings

Below graph shows the coating options that could be used on the spacecraft giving an output with the solarabsorption and emittance

http://www.aticourses.com/sampler/Spacecraft_Thermal_Control.pdf

When choosing the outer surface coating layer it is very important to pick a layer that has the lowest solar absorption (). When looking at the graph above its clear to us that FEP Teflon, with Ag, Al mixture has the lowest absorption but the emittance () variation varies from 0.28- 0.89 which is not acceptable as maintaining a high emittance () is also an important factor when choosing the outer layer of the spacecraft surface coating.

From the above graph it is deduced that the outer surface of this spacecraft (parabolic antenna) will be coated with white paint due to its less solar absorption () that is approximately 0.18 - 0.28. This also has anemittance () of 0.79 – 0.89. As of these properties white paint is the most suitable coating for the outer surface to maintain its survival temperature of 253k-343k. the interior section of the spacecraft will be coated with the coating which has the highest temperature absorption () and the highest emittance (). From the graph the only option, which is clearly acceptable is black paint that has aemittance () rate from 0.83-0.9 and temperature absorption () of 0.92-0.98. Due to these properties of black paint energy enhancement between components will be high giving rise to an equal thermal distribution in between components.

Properties of both black and white paint

density

1000

Kg/m3

thickness

0.0009

m

Surface area

2.5

m2

voloume

0.00225

m3

mass

2.25

kg

Thermal environment analysis

At the launch environment for Arian five launch vehicle, the loading and the moving of the spacecraft has to be maintained within the temperature (284 k – 297 K) which is the tolerable thermal property of the spacecraft. The outside temperature at the launch pad varies between

283 – 298 K. Below table represents spacecraft’s loading, moving and thermal specifications as both chemical and electrical spacecraft will use the Arian five launcher. These specifications are generalized for both the systems.

http://www.scribd.com/doc/16927713/ESAs-Ariane-5-Launch-Vehicle-Users-Manual

In the ariane-5 launch vehicle the maximum accepted aero thermal flux is 1135W/m2the flux is calculated perpendicular to the velocity direction on a plane surface as free molecular flow acting on it.

Albedo, solarradiation flux and terrestrial infraredradiation is added to the aero thermal flux when calculating the incident flux on the space craft altitude of the lounge vehicle, position of the son with respect to the lounge vehicle, orientation and the space craft surface orientation has to been taken in to account for further calculation

For a day light situation with long ballistic and boosted faces radiation of the sun is also been considered in order to reduce the heat flux to lounge the maximum spun is 2deg-s

Above graph represent the aero thermal fluxes on trajectory, the red line represent the total absorb flux. Space craft reaches a maximum flux of 1090 W/m2 and thru out the journey a significant variation of the flux is represented on the graph depending on the time of light the mean absorb flux variation during flight is represented by the black line the entire space craft has to have the ability to withstand a maximum flux of 1090 W/m2for period of approximately 300s without access heating as a solution multi layer insulation (MLI) has been used to control the space craft surface temperature within the torralable range

http://www.scribd.com/doc/16927713/ESAs-Ariane-5-Launch-Vehicle-Users-Manual

Spacecraft Thermal Environment On the earth

Earth’s energy budget percentages

Incoming solar energy 100%

Reflected by atmosphere.6%

Reflected from earth surface.4%

Reflected by clouds.20%

Carried to clouds&atmosphere by latent heat in water vapor 23%

Absorbed by clouds. 3%

Absorbed by atmosphere.

16%

Radiated from clouds &atmosphere. 64%

Radiation absorbed by atmosphere. 15%

Radiated directly to space from earth 6%

Rising air and conduction 7%

Absorbed by land & Ocean 51%

Geostationary transfer orbit

At the time of geostationary transfer the spacecraft is meant to heat to a very high 398 Kbut an acceptable heat where the MLI blankets will reduce the heat to a tolerable value 322 K.

The spacecraft will spend a much higher time in this orbit due the inclination of 12  of the spacecraft, there for exposure to both inner and outer van Allen belts are much higher giving rise to a harsh radiation environment with high eccentricity where the spacecraft will spend a long eclipse period. During the eclipse the spacecraft internal temperature will decrease to about 2 C- 25 C where the radiators and the loop heat pipes system will ensure the inner temperature remains tolerable with in the particular range. The protons and the electrons are the main concerns of the van Allen radiation belts, a representation of proton and electron flux is presented in the bellow graph, the blue and red dotted line is a representation of the spacecraft’s voyage through the radiation belts.

Omnidirectional flux (protons/ cm2-s)

R- plot of constant intensity flux contours with an energy of  1.0 Mev

http://www.braeunig.us/apollo/apollo11-TLI.htm

From the above charts its deduced that the spacecraft requires protection to withstand a maximum of 1.5 x 107Mev

Spacecraft in the Van Allen Radiation Belt region, showing Geosynchronous Transfer Orbit

http://moon-landings.tripod.com/id17.htm

Spacecraft thermal environment as it leaves the earth radius 6378km

Solar Flux

qs = Gs = 1418W/m2 @ winter solstice

= 1326 W/m2 @summer solstice

Emitted

Albedo

qr = a = 30%  5%

http://gsa.confex.com/gsa/2005AM/finalprogram/abstract_95876.ht

Trans lunar injection (TLI)

Trans lunar injection is the propulsive maneuver used to set the spacecraft trajectory for arrival on the lunar orbit. The chemical trans lunar injection last approximately about 133200 s where with time the temperature of the spacecraft decreases due to the decrease in earth’s albedo effect with increasing altitude from earth. At the burn the spacecraft reaches up to a temperature of 320k but with time it settles down gradually to a stable temperature about 287k that is within the spacecraft survival temperature. During the eclipse of the TLI the spacecraft temperature will decrease to a minimum of 236k-240k where the heat pipes will bring back the spacecraft’s thermal environment to the survival range

For the electrical TLI the temperature remained almost stable at about 276k even though at the very short eclipse the temperature varied from 276k-268k keeping the spacecraft at its survival temperature.

Space radiation

Space radiation can be defined as energy packages of particles traveling at a very high speed in space. This can be divided in two parts, ionizing and non ionizing radiation. [1]

• Ionizing radiation–this is radiation that has adequate energy to remove electrons from the orbits of atoms. This results charged particles, and it is this radiation type that is evaluated for requirements ofsufficientsafety. This includes Gamma rays, protons, and neutrons. In ordinary chemical reactions, ionizing radiation forms ions differently, such as sodium and chlorine is a production of table salt and In these kinds of reactions, an ion is formed detaching the outermost electron, which is positively charged.in a ionizing radiation, if there isadequateenergy the electrons other than those in the outermost orbits can be removed.This process makes the atom very unstable giving rise to the ion to be highly chemically reactive.

• Non-ionizing radiation - these are radiation that has no sufficient energy to remove electronsfrom their orbits. These include microwaves, radio waves, and visible light

when considering radiation protection in space, the electromagnetic waves or non-ionizing radiation have little importance .the focus will be onparticle radiation which hasthree naturally occurring sources of particle radiation in space.Trapped radiation, galactic cosmic radiation, and solar particle events could be named as these. [1]

To protect the stagecraft from these radiation adequate shielding should be considered and also when choosing the components of the spacecraft’s, components with good radiation properties must be picked. The structure of the chemical spacecraft will require an aluminum layer of 5mm thick shielding in order to provide approximately 198 mills of protection. The electrical spacecraft will acquire more adequate protection due to its transfer orbit. Therefore an aluminum layer of 6.5mm will be installed to the outer layer that will give protection of approximately 255 mills. The payload section for both space crafts, layer of 7mm aluminum will be used as a shielding due to the sensitive CCD arrays and to protect the sensitive components of the payload.[1] http://radbelts.gsfc.nasa.gov/outreach/space_radiation.html

Thermal environment of the lunar orbit.

Background

2.7 K

Emission

Solar Direct

SC

Sun

5800K

Orbit

Albedo

Emission

NIGHT

DAY

In the lunar orbit the temperature will vary between 287k -310k for both spacecrafts. Due to the moons radiation this high temperature is shown.an occurrence of thermal fatigue showing on both spacecrafts are vulnerable due to the continuously changing heating and cooling as of the frequent eclipse due to the lower allocation of altitude. Radiation around the lunar orbit is very small compared to the radiation exposure due to the voyage. Throughout the life time of the mission the spacecraft will be in a solar minimum. Coronal mass ejections, solar flares and geometric storms are expected slightly throughout the mission; therefore the shielings installed to the spacecrafts have the ability to protect the spacecraft during these solar particle events. This layers of protection installed in the spacecrafts will also provide protection from galactic cosmic rays and lunar albedo. Below is a temperature profile to the lunar orbit.

OPERATING TEMPERATURE RANGES

Spacecraft Internal Units

273k- 313k

Telecommunications

Payload units 263k - 323k

Remote sensing payloads

Optical sensors288k -298k Infrared module 233k-303k

Radiometric units 263k- 323k Radar units 283k–313k

Onboard computer 263k – 323k

Telemetry & command units 263k - 323k

Altitude control

Sun & earth sensors 243k -323k Magnetometer 193k -353k

Electronic units 263k – 328kMomentum &

Gyro package 273k – 323k reaction wheels 268k– 318k

Electrical Power

Batteries (NiH2) 268k – 293k Batteries (NiCd) 273k – 298k

Solar arrays 168k – 383k Power control unit 253k –328k

Propulsion

Propellant, filters, valves, Thrusters 280k-338k

Lines 280k – 328k

Thermal Control

Multilayer insulation (MLI) 113k – 523k Heaters, thermostats

Radiators 178k - 333k heat pipes 238k – 333k

Harness

Spacecraft internal 258k – 328k spacecraft external 173k – 373k

Structure

Nonalignment critical 228k – 338k Alignment critical 291k - 294k

Antennas

Parabolic reflector 113k -368k TT&C 208k – 368k

GPS antenna 178k – 343k

Mechanisms

Pyrotechnics 173k – 153k Electric motors 228k – 353k

Deployment hinge 228k -338k solar array drive assembly

238k-333k

The above table represents the temperature ranges of the different components of spacecraft’s. These have been debuted from an analysis of past lunar missions and with the thermal equations shown below.

Assumptions made for thermal analysis calculations

Solar radiation intensity = 1367 W/m2

Moon radiation intensity= 430 W/m2

Moon albedo= 0.119

Earth radiation intensity= 240 W/m2

Earth albedo=0.359

Boltzmann constant = 5.670 x 10-8 W/m2 -K4

Plank’s constant = 6.626055e-34 Ws2

solar flux =1418 W/m2

Convection

q= hAT

h= Heat transfer coefficient

A= Surface Area

T = Temperaturedifference.

Above equation have been used for calculations during launch after faring separation. Convective heat transfer is used for the pump liquid thermal control system.

Radiation

q = T4

 = emissivity at wave length corresponding to temperature T

 = Stefan- Boltzmann’s constant

T= temperature in Kelvin

q is the heat transfer per unit area where T is the surface temperature and this is used to primary heat transfer mechanisms for the spacecrafts.

Planks equation

Eb= 2hc2* 1.

5ech/kT-1

 = Wavelength

h= Planck’s constant

c= speed of light

k= Boltzmann’s constant

above equation is used at any temperature above absolute zero for all materials that emit thermal (blackbody radiation). The rate of total energy emission and the energy distribution across all wavelengths are a function of absolute temperature T for a perfect black body.

Kirchoff’s Law

Monochromatic emissivity = monochromatic absorptivity

For above emissivity and absorptivity changes with wavelength for given materials. At a given frequency Kirchoff’s law deduces that emissivity and absorptivity will remain the same.

Lower equilibrium temperature is shown for white paint on spacecrafts as it absorbs little energy in the frequencies of the solar spectra. IR frequencies associated with the bodies temperature shows a high emissivity.

View factor

A1 F1-2 = A2. F2-1

F1-2 is the view factor from the surface 1 with area A1 to surface area A2.

A2

A1

Function of relative position orientation, geometry and the size of the two surfaces is the view factor.

Its assumed a defused gray surface condition when considering calculations that is , equal emittance of the particular surface in all directions where emissivity and absorptivity are not strong functions of wavelength.

Radiation Equations

q= AT4 (Into deep space)

q= 12A1F1-2 (T14 – T24) (Between two surfaces)

If one surface is earth assume =1 and T= 250 – 260 K

Precisely q= A(T14 – T24) but T deep space is 4 K,  4

These equations are used to calculate the heat transfer between two surfaces once the view factor is calculated.

Conservation of energy

qin- qout+ qdissipated= dissipated = Eint

t

qincident= qabsorbed+ qreflected+qtransmitted

(for transparent materials with no internal dissipation)

Energy absorbed + energy dissipated – energy emitted = 0 in steady state.

It’s referred to the heat generated by equipment for primary dissipated energy. Energy isreflected, transmitted or absorbed for transparent materials that acquire no internal dissipation.

Steady state temperature of insulated surfaces

qin= GSAp*cos

qout = T4Ar

Gscos



T =

Gs= solar flux =1418 W/m2

Absorbed energy = Solar flux x the area x the absorptivity x cosine of the incident angle

Solar flux is approximated at 1418 W/m2 average around earth. This could vary due to distances from sun and other factors. The area of the temperature to the fourth power is proportional to the emitted energy. The last equation shown above gives the equilibrium temperature of the surface with no internal dissipation where in the denominator  is the emissivity and in the numerator  is the absorptivity.

Space radiators

GS*cos()+ QW/ AR - T4 = 0

QW is the heat to be rejected.

QW is the wasted heat that we intend to get rid of in the radiator where in general radiators are oriented to minimize the incident radiation. Using the above equation , we could determine the temperature of the radiator used for an allocated amount of energy elimination , or a determination of the amount of energy that is removed at a particular temperature from the radiator.

Solar array/flat plate max/ min temperatures

EARTH

sun

H

Energy absorbed includes: from sun on top surface

Qsa= GsAtcos , plus earth IR , plus sunlight reflected from earth.

Earth IR

G1 = q1(4RE2)/(4(H+RE)2)=q1sin2 (energy flux at spacecraft altitude.)

q1=23721 W/m2 (energy flux at at earths surface)

qla=qlsin2Ab (energy absorbed on the bottom surface)

At the surface of the earth it is deduced that the energy flux at a given altitude is lower by the ratio of the area of earth’s surface.to the area of a sphere with radius equals to the altitude of the spacecraft plus the radius of the earth. The bottom equation gives the energy absorbed by the spacecraft.as the radiant energy is infrared emissivity is used in this equation. Choosing + or – is based on on the calculation where you take the maximum or the minimum temperatures.

Solar energy reflected from earth

QAa=GsaAbKasin2

Ka=0.664+0.521+0.2032

a= albedo (percentage of direct solar energy reflected off the earth = 30% 5%

earth reflects the suns energy (albedo) , this is a strong function of the spacecraft and is represented by .

Solar Array/Flat Plate Energy Balance

Absorbed energy

Qa=GsAt+q1Absin2+GsaAbKasin2

Emitted energy Qe=bAT4tAT4

qabsorbed – qemitted – q power generated = 0

qpower generated = GsA where  is the solar array installed efficiency

Both the top and the bottom of the array’s emitted energy are included in the emission, which may have different emissivities. Energy produced by the solar array are subtracted in the energy balance, which incident solar energy is a fraction of it.

Solar array Maximum/Minimum temperatures

1

4

Gst+qlbsin2+GsabKasin2-Gs

A

Tmax =

1

4

qlbsin2

(b+t)

Tmin=

Equilibrium temperature of the array has been rearranged from the equations. Maximum is in the full sun where minimum is for eclipse conditions for functions of absorptivity of the array and its efficiency for a given altitude.

1

4

Spherical Satellite Max/Min tempreatures

AcGs+AFql+AFGsaKa+QW

A

Tmax s =

1

4

AFql+QW

A

Tmin s =

A = satellite surface area, Ac is the satellite cross – sectional area , and F is the view factor = (1-cos )/2

Multi layer insulation

Modeled at an effective emissivity

q= effective(Th4- Tc4)

Thermal conclusion

MLI mass

4.4484

Kg

Additional thermal controller mass (heat pipes ect)

0.8616

Kg

Radiators mass

2x4.32

Kg

Paint mass

2.25

Kg

Total mass

16.2

Kg

Heat pipes power

5

W

Total radiator power

2x2

W

Total power

9

W

Above table represents the final mass and power allocation for both chemical and electrical options. Even though the chemical option showed a slightly higher surface area due to its fuel tanks the passive thermal protection requirements have been taken into consideration generally due to the loop heat pipes multiple evaporators and the advanced variable emissivity radiator systems gives n higher thermal protection system to maintain a equipment surviving temperature within both space crafts throughout the voyage both space crafts will maintain an internal temperature of 278k-298k. For the above final mass and power allocation information has been taken from other subsystems about the engine design, surface areas, volume and the power consumption of the overall spacecraft. Initially iteration started with the mass allocation 25kg and a maximum power allocation 18W for both chemical and electrical options. Designing two different thermal controlling subsystems was the most acceptable option due to strict mass constrain of the chemical spacecraft. However when iterations were done lower mass chemical mass budget and high electrical mass budget was shown with slightly approximately equal power budget.

However after calculating the survival temperature ranges, significant thermal temperature changes were not shown for both options even though the trans Luna injection for both space crafts were different, where the electrical space crafts required a higher thermal demand. Higher mass allocation for the electrical spacecraft was not a problem due to the overall mass therefore designing a better thermal protection system that will acquire more mass could be easily implemented for a better spacecraft lifespan with lower risk of failure.

After it was decided that a generalized passive thermal protection system could be made for both spacecrafts. The new loop heat pipes system will give adequate protection higher than the required survival temperatures of internal and external components of both space crafts. As shown in the above table a total mass of 16.2kg and a total power of 9W thermal protection system, is a low powered and a low mass thermal subsystem for the Luna mission


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