Thermal Summary - MAELabs UCSD

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Nov 12, 2013 (3 years and 8 months ago)

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1


Aerospace Thermal Analysis
Overview

G. Nacouzi

ME 155B

2

Thermal Analysis Summary
Overview


Purpose


Summary of heat transfer modes


Aerospace considerations: LV & SC


Thermal control


SC equilibrium temperature


Example



3

Thermal Considerations


Thermal control system for launch and space
vehicles needed to ensure vehicles survive
environment, operate properly and meet customer
requirements.


Material properties, ultimate strength are a
function of temperature


Temperature gradients can cause pointing
errors in space vehicles, and stresses on
structure (Payload cooling, e.g., cryostat not
addressed)

4

Thermal Analysis Summary
Review


Quick review of the 3 modes of heat transfer:


Conduction, driven by material thermal
conductivity k: q ~ k dT/dx


Convective (Forced and natural), driven by
convective heat transfer coefficient, h:


q ~ hA(

T)


Radiative, driven by surface properties & view
factor F: q ~

F T
4



where


is the surface emissivity coefficient

5

Aerospace Considerations


Pre
-
launch environmental conditions: easily
controlled


Launch phase conditions


Launch vehicle thermal envt: Dominated by
aeroheating (forced convection), other
contributors include combustion chamber,
plume and limited solar


SV may be subjected to limited, usually
insignificant, heating envt during launch phase

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Aerospace Considerations


On
-
orbit thermal environment


SV environment dominated by radiative
(internal & external) & conduction from
internally generated sources. Significant envt
difference on SC, e.g., Sun heating and deep
space cooling. Some molecular heating may
occur depending on orbit.


Re
-
entry heating dominated by convective and
radiative => requires ablative or high
temperature, e.g., ceramics, materials

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Launch Vehicle Considerations


Convective heat transfer calculations based
on: Vehicle drag, i.e., skin friction
coefficient, shock structures, flow regime
and transition points


Trajectory parameters, i.e., flight time,
velocity & altitude profiles (q ~

V
3
)


Internal heat sources, i.e., combustion
chamber

8

Launch Vehicle Considerations


Several light weight insulators are available
to protect launch vehicle and maintain
structure and other components below
maximum allowable temperature.


Material uses include different coatings, low
density ablative materials and light weight
insulators


High density ablators such as carbon
-
carbon
and carbon phenolics are used for hot nozzle
operation

9

Space Vehicle Thermal
Considerations


Thermal envt dominated by radiative transfer
and internally generated heat


Main external contributors:


Emission from Sun, Js. Sun assumed as a blackbody
with a Temp. of 5800 K, usually a point source.

Js = (P/ 4 pi a*a), where,

P ~ Sun total emission = 3.856 E26 W

a ~ distance from planet to Sun => Js = 1370 W/m^2
for Earth


Radiation into deep space, assume 0 K

10

Space Vehicle Thermal
Considerations


Main external contributors (cont’d):


Emission from Earth, effective Earth Temp ~ 290 K =>
Earth emits in the long IR region between 2 & 50 um


Wien’s law
-
> max (

T) ~ 2890 um K


Reflection from Earth. Earth reflects some of the
impinged Sun energy back into space. Fraction based
on Albedo, w. Effective w ranges from 0.3 to 0.4
(instantaneous w has wider range, depending on
reflecting ‘surface’)

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Space Vehicle Thermal Control


Active
: Requires control (& power) from S/C


Includes: Louvers (shutters), heaters, coolers,
cryo. Systems


Passive
: Self contained system, no power
input. Includes:


Geometry & structure, insulators, shields,
radiating fins, phase change materials and heat
pipes

12

Space Vehicle Thermal Control:
Heat Pipes


Heat pipe contains a wick running the length
of the sealed pipe which is partially filled with
a fluid such as ammonia.


One end of the heat pipe is connected to a hot
region, while the other end is exposed to the
colder portion.


Heat causes fluid to evaporate (at hot end) and
condense at cold end. Wick transports fluid back
to hot end through capillary motion. Heat of
vaporization used to cool hot end.

13

Space Vehicle Thermal Control:
Heat Pipes

Evaporation

Heat in

Vapor

Liquid

Condensation

Heat

out

Wick

Heat transfer based on latent heat of vaporization

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SC Thermal Balance

Jrad



T
4


Jabs

Jin


For equilibrium condition:
Jabs
Aabs

Jrad
Arad


T
4
Aabs
Arad
Jin












Jrad ~ radiated heat flux

Jabs ~ absorbed heat flux

A ~ Emissive or abs. Area


,


~ absorption and emissive


coefficient

Note that there is a spectral dependency for

,

. For the Sun (

), peak

is at about 0.45 um while for Earth peak is at ~ 10 um

(Wien’s law)

15

SC Thermal Balance: Example

Ja

Jearth

Js

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SC Thermal Balance: Example


Heat from Sun: Js x Asun x



Sun reflected from Earth: Ja x


x Aalb


Emission from Earth: Jearth *


* Aearth


Heat radiated to space:



Aspace * T
4


Internal heat generation: Q

Heat balance calculation:


(Asun Js + Aalb Ja)*


+ Aearth Jearth *




+ Q =



Aspace * T
4


17

SC Thermal Balance: Example

T
4

= Aearth * Jearth/(Asun

) + Q/(Asun



)


+ (Asun Jsun + Aalb Ja)/(Asun

) (

/

)

where, Area (Ax) refers to x surface involved in
process, e.g., Aearth is the area of the SC that is
exposed to Earth.



Approach provides average temperature of
SC, used only as initial estimate
.

18

SC Thermal Analysis


Example presented a ‘zeroth’ order analysis
to estimate equilibrium temperature


Actual temperature distribution solved
using a nodal network representing SC and
accounts for all modes of heat transfer at
each node.


Needed to determine local hot spots and
temperature gradients