ASTM E491-73(2020)

This international standard was developed in accordance with internationally recognized principles on standardization established in the Decision on Principles for the
Development of International Standards, Guides and Recommendations issued by the World Trade Organization Technical Barriers to Trade (TBT) Committee.
Designation: E491 − 73 (Reapproved 2020)
Standard Practice for
Solar Simulation for Thermal Balance Testing of Spacecraft
This standard is issued under the fixed designation E491; the number immediately following the designation indicates the year of
original adoption or, in the case of revision, the year of last revision.Anumber in parentheses indicates the year of last reapproval.A
superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope 1.3.2 The ultimate result of the thermal balance test is to
prove the thermal design to the satisfaction of the thermal
1.1 Purpose:
designers. Flexibility must be provided to them to trade off
1.1.1 The primary purpose of this practice is to provide
additional analytical effort for simulator shortcomings. The
guidance for making adequate thermal balance tests of space-
combination of a comprehensive thermal-analytical model,
craft and components where solar simulation has been deter-
modern computers, and a competent team of analysts greatly
mined to be the applicable method. Careful adherence to this
reduces the requirements for accuracy of space simulation.
practice should ensure the adequate simulation of the radiation
environment of space for thermal tests of space vehicles. 1.4 Utility—This practice will be useful during space ve-
1.1.2 A corollary purpose is to provide the proper test hicle test phases from the development through flight accep-
environmentforsystems-integrationtestsofspacevehicles.An tance test. It should provide guidance for space simulation
accurate space-simulation test for thermal balance generally testing early in the design phase of thermal control models of
willprovideagoodenvironmentforoperatingallelectricaland subsystems and spacecraft. Flight spacecraft frequently are
mechanical systems in their various mission modes to deter- tested before launch. Occasionally, tests are made in a space
mine interferences within the complete system. Although chamber after a sister spacecraft is launched as an aid in
adherence to this practice will provide the correct thermal analyzing anomalies that occur in space.
environment for this type of test, there is no discussion of the
1.5 This standard does not purport to address all of the
extensive electronic equipment and procedures required to
safety concerns, if any, associated with its use. It is the
support systems-integration testing.
responsibility of the user of this standard to establish appro-
priate safety, health, and environmental practices and deter-
1.2 Nonapplicability—This practice does not apply to or
mine the applicability of regulatory limitations prior to use.
provide incomplete coverage of the following types of tests:
1.6 This international standard was developed in accor-
1.2.1 Launch phase or atmospheric reentry of space
dance with internationally recognized principles on standard-
vehicles,
ization established in the Decision on Principles for the
1.2.2 Landers on planet surfaces,
Development of International Standards, Guides and Recom-
1.2.3 Degradation of thermal coatings,
mendations issued by the World Trade Organization Technical
1.2.4 Increased friction in space of mechanical devices,
Barriers to Trade (TBT) Committee.
sometimes called “cold welding,”
1.2.5 Sun sensors,
2. Referenced Documents
1.2.6 Man in space,
1.2.7 Energy conversion devices, and
2.1 ASTM Standards:
1.2.8 Tests of components for leaks, outgassing, radiation E259Practice for Preparation of Pressed Powder White
damage, or bulk thermal properties.
Reflectance Factor Transfer Standards for Hemispherical
and Bi-Directional Geometries
1.3 Range of Application:
E296Practice for Ionization Gage Application to Space
1.3.1 The extreme diversification of space-craft, design
Simulators
philosophies, and analytical effort makes the preparation of a
E297Test Method for Calibrating Ionization Vacuum Gage
brief, concise document impossible. Because of this, various
Tubes (Withdrawn 1983)
spacecraftparametersareclassifiedandrelatedtotheimportant
E349Terminology Relating to Space Simulation
characteristic of space simulators in a chart in 7.6.
1 2
This practice is under the jurisdiction of ASTM Committee E21 on Space For referenced ASTM standards, visit the ASTM website, www.astm.org, or
Simulation andApplications of SpaceTechnology and is the direct responsibility of contact ASTM Customer Service at service@astm.org. For Annual Book of ASTM
Subcommittee E21.04 on Space Simulation Test Methods. Standards volume information, refer to the standard’s Document Summary page on
Current edition approved Nov. 1, 2020. Published December 2020. Originally the ASTM website.
approvedin1973.Lastpreviouseditionapprovedin2015asE491–73(2015).DOI: The last approved version of this historical standard is referenced on
10.1520/E0491-73R20. www.astm.org.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
E491 − 73 (2020)
2.2 ISO Standard: 3.2.5 albedo—the ratio of the amount of electromagnetic
ISO 1000-1973SI Units and Recommendations for the Use radiation reflected by a body to the amount incident upon it.
of Their Multiples and of Certain Other Units
3.2.6 apparent source—the minimum area of the final ele-
2.3 American National Standards:
ments of the solar optical system from which issues 95% or
ANSI Y10.18-1967Letter Symbols for Illuminating Engi-
more of the energy that strikes an arbitrary point on the test
neering
specimen.
ANSIZ7.1-1967StandardNomenclatureandDefinitionsfor
3.2.7 astronomical unit (AU)—a unit of length defined as
Illuminating Engineering
the mean distance from the earth to the sun (that is,
ANSI Y10.19-1969Letter Symbols for Units Used in Sci-
149597890 6 500 km).
ence and Technology
3.2.8 blackbody (USA),Planckian radiator—a thermal ra-
diator which completely absorbs all incident radiation, what-
3. Terminology
ever the wavelength, the direction of incidence, or the polar-
3.1 Definitions, Symbols, Units, and Constants—This sec-
ization. This radiator has, for any wavelength, the maximum
tioncontainstherecommendeddefinitions,symbols,units,and
spectral concentration of radiant exitance at a given tempera-
constantsforuseinsolarsimulationforthermalbalancetesting
ture (E349).
of spacecraft. The International System of Units (SI) and
3.2.9 collimate—to render parallel, (for example, rays of
International and American National Standards have been
light).
adheredtoasmuchaspossible.TerminologyE349isalsoused
3.2.10 collimation angle—in solar simulation, the angular
andissoindicatedinthetext.Table1providescommonlyused
nonparallelism of the solar beam, that is, the decollimation
symbols.
angle. In general, a collimated solar simulator uses an optical
3.2 Definitions:
componenttoimageatinfinityanapparentsource(pseudosun)
3.2.1 absorptance (α , α ,α )—ratio of the absorbed radiant
e v
offinitesize.Theanglesubtendedbytheapparentsourcetothe
or luminous flux to the incident flux (E349)(Table 1).
finalopticalcomponentreferredtoasthecollimator,isdefined
3.2.2 absorptivity of an absorbing material—internal ab-
asthesolarsubtenseangleandestablishesthenominalangleof
sorptance of a layer of the material such that the path of the
decollimation. A primary property of the “collimated” system
radiation is of unit length (E349).
is the near constancy of the angular subtense angle as viewed
3.2.3 air mass one (AM1)—the equivalent atmospheric at- from any point in the test volume. The solar subtense angle is
tenuation of the electromagnetic spectrum to modify the solar therefore a measure of the nonparallelism of the beam. To
irradiance as measured at one astronomical unit from the sum avoid confusion between various scientific fields, the use of
outside the sensible atmosphere to that received at sea level, solar subtense angle instead of collimation angle or decollima-
when the sun is in the zenith position. tion angle is encouraged (see solar subtense angle).
3.2.4 air mass zero (AM0)—the absence of atmospheric 3.2.11 collimator—an optical device which renders rays of
attenuation of the solar irradiance at one astronomical unit light parallel.
from the sun.
3.2.12 decollimation angle—not recommended (see colli-
mation angle).
3.2.13 diffuse reflector—a body that reflects radiant energy
in such a manner that the reflected energy may be treated as if
Withdrawn.
it were being emitted (radiated) in accordance with Lambert’s
Available fromAmerican National Standards Institute (ANSI), 25 W. 43rd St.,
4th Floor, New York, NY 10036, http://www.ansi.org.
TABLE 1 Commonly Used Symbols
Symbol Quantity Definition Equation or Value Unit Unit Symbol
Q radiant energy, work, joule J
quantity of heat
−1
Φ radiant flux Φ =dQ/dt watt (joule/second) W, Js
−2
E irradiance (receiver) flux E =dΦ/dA watt per square metre W·m
density
−2
M radiant exitance (source) M =dΦ/dA watt per square metre W·m
−1
I radiant intensity (source) I =dΦ/dω watt per steradian W·sr
ω = solid angle through which flux from source is radiated
−1 −2
L radiance L =dI/(dA cosθ ) watt per steradian = W·sr ·m
square metre
θ = angle between line of sight and normal to surface dA
τ transmittance τ = Φ, transmitted/Φ, incident none
τ(λ) spectral transmittance τ(λ)= Φ(λ), transmitted/Φ(λ), incident none
ρ reflectance (total) ρ = Φ, reflected/Φ, incident none
εH emittance (total εH = M, specimen/M, blackbody
hemispherical)
α absorptance α = Φ, absorbed/Φ, incident none
α solar absorptance α = solar irradiance absorbed/solar irradiance incident none
s s
E491 − 73 (2020)
law.The radiant intensity reflected in any direction from a unit 3.2.27 integrating (Ulbrecht) sphere—part of an integrating
area of such a reflector varies as the cosine of the angle photometer. It is a sphere which is coated internally with a
between the normal to the surface and the direction of the white diffusing paint as nonselective as possible, and which is
reflected radiant energy (E349). provided with associated equipment for making a photometric
measurement at a point of the inner surface of the sphere. A
3.2.14 dispersion function (X/λ)—a measure of the separa-
screen placed inside the sphere prevents the point under
tion of wavelengths from each other at the exit slit of the
observation from receiving any radiation directly from the
monochromator, where X is the distance in the slit plane and λ
source (E349).
is wavelength. The dispersion function is, in general, different
3.2.28 intensity—see radiant intensity.
for each monochromator design and is usually available from
the manufacturer.
3.2.29 irradiance at a point on a surface E ,E;E =dΦ /
e e e
dA—quotient of the radiant flux incident on an element of the
3.2.15 divergence angle—see solar beam divergence
surface containing the point, by the area of that element
angle(3.2.60).
−2
measured in W·m (E349)(Table 1).
3.2.16 electromagnetic spectrum—the ordered array of
¯
3.2.30 irradiance, mean total (E)—the average total irradi-
known electromagnetic radiations, extending from the shortest
ance over the test volume, as defined by the following
wavelengths, gamma rays, through X rays, ultraviolet
equation:
radiation, visible radiation, infrared and including microwave
and all other wavelengths of radio energy (E349).
¯
E 5 E r,θ,z dV/ dV (1)
* ~ ! *
v v
3.2.17 emissivity of a thermal radiator ε, ε =M /
e,th
where:
M (ε = 1)—ratio of the thermal radiant exitance of the radiator
e
¯
E(r,θ,z) = total irradiance as a function of position (Table
to that of a full radiator at the same temperature, formerly
1).
“pouvoir emissif” (E349).
3.2.31 irradiance, spectral [E or E(λ)] —the irradiance at a
3.2.18 emittance (ε)—the ratio of the radiant exitance of a λ
specific wavelength over a narrow bandwidth, or as a function
specimen to that emitted by a blackbody radiator at the same
of wavelength.
temperature identically viewed. The term generally refers to a
specific sample or measurement of a specific sample. Total
3.2.32 irradiance, temporal—the temporal variation of in-
hemispherical emittance is the energy emitted over the hemi- dividual irradiances from the mean irradiance. The temporal
sphere above emitting element for all wavelengths. Normal
variations should be measured over time intervals equal to the
emittance refers to the emittance normal to the surface to the
thermal time constants of the components. The temporal
emitting body.
stability of total irradiance can be defined as:
3.2.19 exitance at a point on a surface (radiant exitance) ¯
E 56100@ ∆E 1∆E /2E# (2)
~ !
t t ~min! t ~max!
(M)—quotient of the radiant flux leaving an element of the
3.2.33 irradiance, total—the integration over all wave-
surface containing the point, by the area of that element,
−2 lengths of the spectral irradiance.
measured in W·m (E349)(Table 1).
3.2.34 irradiance, uniformity of—uniformity of total irradi-
3.2.20 field angle—not recommended (see solar beam sub-
ance can be defined as:
tense angle).
¯
E 56100@ E 1E /2E# (3)
~ !
u ~min! ~max!
3.2.21 flight model—an operational flight-capable space-
craft that is usually subjected to acceptance tests.
where:
3.2.22 flux (radiant, particulate, and so forth)—for electro-
E = uniformity of the irradiance within the test volume,
u
magnetic radiation, the quantity of radiant energy flowing per
expressed as a percent of the mean irradiance,
unit time; for particles and photons, the number of particles or E = smallest value obtained for irradiance within the
(min)
photons flowing per unit time (E349). test volume, and
E = largest value obtained for irradiance within the test
(max)
3.2.23 gray body—a body for which the spectral emittance
volume.
and absorptance is constant and independent of wavelength.
Uniformity of irradiance values must always be specified
The term is also used to describe bodies whose spectral
together with the largest linear dimension of the detector used.
emittance and absorptance are constant within a given wave-
length band of interest (E349).
3.2.35 Lambert’s law—the radiant intensity (flux per unit
solid angle) emitted in any direction from a unit-radiating
3.2.24 incident angle—the angle at which a ray of energy
surface varies as the cosine of the angle between the normal to
impinges upon a surface, usually measured between the direc-
the surface and the direction of the radiation (also called
tion of propagation of the energy and a perpendicular to the
Lambert’scosinelaw).Lambert’slawisnotobeyedexactlyby
surface at the point of impingement or incidence.
most real surfaces, but an ideal blackbody emits according to
3.2.25 infrared radiation—see electromagnetic spectrum
this law. This law is also satisfied (by definition) by the
(E349).
distributionofradiationfromaperfectlydiffuseradiatorandby
3.2.26 insolation—direct solar irradiance received at a the radiation reflected by a perfectly diffuse reflector. In
surface, contracted from incoming solar radiation. accordance with Lambert’s law, an incandescent spherical
E491 − 73 (2020)
blackbody when viewed from a distance appears to be a
πhc = c /2 (for the exitance of the polarized radiation)
uniformlyilluminateddisk.Thislawdoesnottakeintoaccount
2hc = c /π (for the radiance of the nonpolarized radiation)
any effects that may alter the radiation after it leaves the hc = c /2π (for the radiance of the polarized radiation)
8πhc =4c /c (for the energy per unit volume of the nonpo-
source.
larized radiation
3.2.36 maximum test plane divergence angle—the angle
3.2.40 prototype model—a spacecraft or subsystem that is
between the extreme ray from the apparent source and the test
used for development or qualification test. This is an accurate
plane. This applies principally to direct projection beams
reproductionofactualspacehardwareandisidenticalornearly
whereitisequivalenttoonehalftheprojectionconeangle(see
identical to the flight model.
Fig. 1).
3.2.41 pyranometer—an instrument that measures the com-
3.2.37 natural bandwidth—the width at half height of a
binedsolarirradianceanddiffuseskyirradiance.Thepyranom-
radiationsourceemissionpeak.Itisindependentofinstrument
eter consists of a recorder and a radiation-sensing element
spectral bandwidth, being an intrinsic property of the radiation
which is mounted so that it views the entire sky.
source.
3.2.42 pyrheliometer—an instrument that measures the di-
3.2.38 penumbra—see umbra.
rect solar irradiance, consisting of a casing which is closed
except for a small aperture through which the direct solar rays
3.2.39 Planck’s law—alawgivingthespectralconcentration
enter, and a recorder unit.
ofradiantexitanceofafullradiatorasafunctionofwavelength
3.2.43 Angstrom compensation pyrheliometer—an instru-
and temperature. For the total radiation emitted (unpolarized):
mentdevelopedbyK.Angstromforthemeasurementofdirect
25 c2λT 21
M λ,T 5 c λ e 21 (4)
~ ! ~ !
solar irradiation. The radiation receiver station consists of two
where:
identical manganin strips whose temperatures are measured by
−2
attached thermocouples. One of the strips is shaded, whereas
M = spectral concentration, W·m ;
λ = wavelength, m; and
theotherisexposedtosunlight.Anelectricalheatingcurrentis
T = absolute temperature, K.
passedthroughtheshadedstripsoastoraiseitstemperatureto
that of the exposed strip. The electric power required to
The constants are:
accomplish this is a measure of the solar irradiance.
2 216 22
c 52π hc 53.741844 310 W·m (5)
3.2.44 radiance (in a given direction, at a point on the
c 5 hc/k 51.438833 310 m·K surface of a source or receptor, or at a point in the path of a
beam)—quotient of the radiant flux leaving, arriving at, or
where:
passing through an element or surface at this point, and
h = Planck’s constant,
propagated in directions defined by an elementary cone con-
c = velocity of light in vacuum, and
taining the given direction by the product of the solid angle of
k = Boltzmann constant.
the cone, and the area of the orthogonal projection of the
NOTE1—Itisrecommendedthattheconstant c isalwaysusedwiththe
element of surface on a plane perpendicular to the given
meaningnotedabove.Thenumericalconstantsapplicabletootheraspects
direction (E349)(Table 1). Symbol: L , L; L =d Φ/(dωdA cos
e e
of the radiation emitted are shown below. They should be designated c
−1 −2
θ); measured in W·sr m .
multiplied by an appropriate factor.
3.2.45 radiant flux (φ)—radiant power, power-emitted,
transferred, or received as radiation, measured in W (E349)
(Table 1).
3.2.46 radiant flux (surface)density at a point of a surface—
quotient of the radiant flux at an element of the surface
containing the point, by the area of that element (also see
−2
irradiance and radiant exitance), measured in W·m (E349).
3.2.47 radiant intensity of a source, in a given direction
(I)—quotient of the radiant flux leaving the source propagated
in an element of solid angle containing the given direction, by
−1
the element of solid angle measured in W · sr (E349)(Table
1).
NOTE 2—For a source that is not a point source: The quotient of the
radiantfluxreceivedatanelementarysurfacebythesolidanglewhichthis
surface subtends at any point of the source, when this quotient is taken to
the limit as the distance between the surface and the source is increased.
3.2.48 radiation, monochromatic—radiation at a single
wavelength, and by extension, radiation of a very small range
of frequencies or wavelengths.
FIG. 1 Solar Subtense and Divergence Angles NOTE3—Useoftheadjective“spectral.”Whencertainproperties,such
E491 − 73 (2020)
as absorptance or transmittance, and so forth, are considered for mono-
3.2.59 solar absorptance (α )—the ratio of the absorbed
s
chromaticradiation,andtheyarefunctionsofwavelength(orfrequencyor
solar flux to the incident solar flux (Table 1).
wavenumber, and so forth), the term may be preceded by the adjective
` `
“spectral” or by the property symbol followed by the subscript λ, or both;
α 5 * α~λ!E~λ!dλ/* E~λ!dλ (6)
0 0
example: spectral transmittance τ(λ)(E349).
3.2.60 solar beam divergence angle—the angle measured
3.2.49 radiometer—instrument for measuring irradiance in
fromalineextendingfromthecenteroftheapparentsourceto
energy or power units (E349).
anarbitrarypointinthetestvolumeandtoalineparalleltothe
3.2.50 radiometry—measurement of the quantities associ-
principal axis of the solar beam (see Fig. 1).
ated with irradiance (E349).
3.2.61 solar beam incident angle—theanglemeasuredfrom
3.2.51 reflection—return of radiation by a surface without
a line extending from the center of the apparent source to an
change frequency of the monochromatic components of which
arbitrary point on the test specimen and the surface normal at
the radiation is composed (E349).
that point.
3.2.52 reflection, diffuse—reflection in which, on the micro-
3.2.62 solar beam subtense angle—that angle subtended by
scopic scale, there is no specular reflection (E349).
the maximum dimension of the apparent source at an arbitrary
3.2.53 reflection, mixed—partly specular and partly diffuse-
point on the test specimen (see Fig. 1).
reflected (E349).
NOTE 4—The terms “collimation angle” and “field angle” are some-
3.2.54 regular (specular)reflection—reflection without dif-
times used for “subtense angle.” The term “subtense angle” is preferred.
fusioninaccordancewiththelawsofopticalreflection(E349).
3.2.63 solar constant—the total solar irradiance at normal
3.2.55 resolution—a qualitative term relating to the fidelity
incidence on a surface in free space at the earth’s mean
of reproduction of the natural band (both in height and width).
distance from the sun (1 AU).
An emission peak is said to be completely resolved when the
−2
NOTE5—Thecurrentacceptedvalueof1AUis1353 621W·m and
observed band is practically identical to the natural band. Fig.
is subject to change.
2 shows the relationship between resolution (observed peak
3.2.64 space environment simulation—a laboratory duplica-
height/true peak height) and the ratio of spectral bandwidth to
tion of one or more of the effects of the space environmental
natural bandwidth. Note that when this ratio is small, the
parameters on a spacecraft, components, or materials. The
deviation from true peak height is small, the fraction being
natural environmental parameters include vacuum-pressure,
99.6% at a ratio of 0.1.
particulate radiation, electromagnetic radiation, and meteroid
3.2.56 reflectance (ρ)—ratio of the reflected radiant or
radiation.Inducedenvironmentalparametersincludevibration,
luminous flux to the incident flux (E349)(Table 1).
shock, and acceleration. The effects can include thermal
3.2.57 reflectivity—reflectanceofalayerofmaterialofsuch
balance, heat transfer, material property change, operational/
athicknessthatthereisnochangeofreflectancewithincreased mechanical subsystem problem, and subsystem functional
thickness (E349).
testing.
3.2.58 slit width—the physical width of a monochromator
3.2.65 spectra, line—the spontaneous emission of electro-
slit opening. In general, all slits should be equal in width at all
magneticradiationfromtheboundelectronsastheyjumpfrom
times. The exit defines the wavelength bandwidth directed to
high to low energy levels in an atom. This radiation is
thedetector.Theenergyincidentuponthedetectorvariesasthe
essentially at a single frequency determined by the jump in
square of the slit width.
energy. Each different jump in energy level, therefore, has its
own frequency and the net radiation is referred to as the line
spectra. Since these line spectra are characteristic of the atom,
they can be used for identification purposes.
3.2.66 spectropyrheliometer—an instrument that measures
the spectral distribution of direct solar irradiance.
3.2.67 spectroradiometer—an instrument for measuring the
spectral concentration of radiant energy or radiant power, also
called “spectrometer” (E349).
3.2.68 spectrum, continuous—a spectrum in which
wavelengths, wavenumbers, and frequencies are represented
by the continuum of real numbers or a portion rather than by a
discrete sequence of numbers (see spectra). For electromag-
netic radiation, it is a spectrum that exhibits no detailed
structure and represents a gradual variation of intensity 0 with
wavelength from one end to the other, such as the spectrum
from an incandescent solid.
3.2.69 spectral filter—an optical component that is spec-
trally selective, or any optical component that rejects radiation
FIG. 2 Relationship of Peak Height to Spectral Bandwidth/Natural
Bandwidth Ratio in spectral regions to shape the resulting spectral distribution.
E491 − 73 (2020)
3.2.70 Stefan-Boltzmann law—the relation between the ra- 4. Summary of Practice
diant exitance of a blackbody radiator and its temperature.
4.1 Thermal balance testing of spacecraft can be performed
M 5σT (7) inmanyways.Thespecificmethodsdependuponsuchitemsas
the spacecraft design, the characteristics of the available
where the constant of proportionality (σ) is called the
simulator,themissionrequirements,thecost,andtheschedule.
Stefan-Boltzmann constant and has a value of
Therefore, it is not desirable or possible to include all thermal
−8 −2 −4
5.66961×10 W·m K .
balance tests in one test method.
3.2.71 subtense angle—see solar beam subtense angle.
4.2 This practice defines terms, discusses test requirements
3.2.72 test volume, simulator—the total volume within the
and instrumentation, and reviews general procedures, safety,
space environmental chamber that can simulate the desired
and maintenance. The test, instrumentation, and thermal engi-
effects.
neers must provide the detailed test method that will satisfy
3.2.73 test volume, spacecraft—the volume occupied by the
their particular requirements and they must be fully aware of
spacecraft within the space simulation chamber throughout the
the effects of the necessary deviations from the ideal.
duration of the test. Unless otherwise specified, test volume is
5. General Considerations
meant to mean spacecraft test volume.
3.2.74 thermal analytical model—a mathematical model of 5.1 The use of solar simulation for thermal balance testing
of spacecraft imposes a number of specific technical require-
thethermalcharacteristicsofaspacecraftthatisusuallysolved
using a computer. ments and methods. The general considerations covered here
relate more to the philosophical bases of the various thermal
3.2.75 thermal balance test—a test or series of tests con-
balance tests rather than to their specific implementation.
ducted upon a spacecraft or model to determine the tempera-
tures in space under normal or extreme operating conditions. 5.2 A space program can be said to have its own unique
Both transient and equilibrium conditions can be simulated. characteristics and problems and the same can be said for each
test facility. The characteristics of both the facility and the test
3.2.76 thermal radiator—source-emitting by thermal radia-
itemmustbeconsideredinthedefinitionofthethermalbalance
tion (E349).
tests. First, however, one must establish the purpose of the test
3.2.77 thermopile—a transducer for converting thermal en-
and determine what must be proved or verified. Second, one
ergy directly into electrical energy, composed of pairs of
may devise an excellent test program assuming no monetary,
thermocouples which are connected either in series or in
schedule or facility limitations. Finally, one may recognize the
parallel.
restraints and establish a set of meaningful compromises.
3.2.78 transmission—passage of radiation through a me-
5.3 This section is separated into four parts:
dium without change of frequency of the monochromatic
5.3.1 Purposes or reasons for performing thermal balance
components of which the radiation is composed (E349).
tests. Each test rationale is related to a specific model of the
3.2.79 transmittance (τ)—ratio of the transmitted radiant
spacecraft; that is, the thermal control model, the qualification
flux to the incident flux (E349)(Table 1).
model, or prototype, and the acceptance or flight model. On
3.2.80 ultraviolet radiation—see electromagnetic spectrum
each of these the test is performed for a slightly different
(E349).
reason.
5.3.2 Ideal Thermal Balance Test Program—This is the
3.2.81 umbra—thedarkestpartofashadowinwhichlightis
program that would be performed if there were no restraints,
completely cut off by an intervening object. A lighter part
suchascost,schedule,andfacilitylimitations.Thisidealtestis
surrounding the umbra, in which the light is only partly cutoff,
also described in terms of thermal control model, prototype
is called penumbra.
model, and flight model spacecraft.
3.2.82 visible radiation—see electromagnetic spectrum
5.3.3 Tradeoff considerations that should be examined be-
(E349).
fore establishing the final test program, and typical test
3.3 Commonly Used Constants—The values of the physical
configurations.
constants presented below are taken from Refs (1) and (2).
5.3.4 Definition and content of the selected program.
The constants are subject to change and the latest available
5.4 Purpose of Thermal Balance Testing—The severity of
supplied by the National Bureau of Standards should be used.
the space thermal environment demands a thorough verifica-
Symbol Constant Value
tion of the thermal design of the spacecraft and its subsystems.
8 −1
To do this, a number of spacecraft models are tested within a
c velocity of light in vacuum 2.997 925·10 m·s
−34
h Planck’s constant 6.626 196·10 J·s
given program. Usually these include a thermal control model,
−16 2
c first radiation constant 3.741 844·10 W· m
aprototype,andoneormoreflightmodels.Ineachofthesetest
−2
c second radiation constant 1.438 833·10 m·K
−3
exposures there are specific, but slightly different reasons, for
b Wien displacement constant 2.899 78 × 10 m·K
−8 −2 −4
σ Stefan-Boltzmann constant 5.669 61 × 10 W·m ·K
performing the test.
5.4.1 Thermal Control Model (Development Test)—Thepur-
pose of the thermal balance test of the thermal model is to
obtain empirical data relating to the spacecraft thermal prop-
Theboldfacenumbersinparenthesesrefertothelistofreferencesattheendof
this practice. erties.Thesedataareintheformoftemperaturemeasurements
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providedbytemperaturetransducersdistributedthroughoutthe significant characteristics. A prime purpose of this test is
spacecraft.Insomecases,asmanyasseveralhundredlocations frequently the verification of the thermal analytical model.
are monitored. During the test exposure various spacecraft Often arbitrary test conditions may be more accurately con-
operational modes may be simulated as well as external trolled and more reproducibly established than the true space
thermal inputs from solar, earth, and lunar simulators.The test environment can be simulated. These known thermal inputs
itemnormallyhasdummyelectronicassemblieswhichprovide maythenbeinsertedasforcingfunctionsforacomputerrunof
a simulation of the mass and thermal dissipation of the actual the analytical model, thus providing a basis for the prediction
units. Both passive and active thermal control techniques are of in-chamber temperatures. The success of these predictions
testedinthismanner.Thedataderivedfromthethermalcontrol establishes the validity of the analytical model. The arbitrary
model test may be used to refine the mathematical model, if test-condition exposures need not replace an accurate orbital
one exists, or may be used directly by the thermal analyst to simulation, but often are performed in addition to it. The ideal
assess the adequacy of the thermal design. thermalcontrolmodeltestconditionsshouldhavenounknown
thermalinputs.Amongthethingsthatshouldbeknownarethe
5.4.2 Prototype (Qualification Test)—The configuration of
differences between the solar simulator and the real in-space
thespacecraftusedforqualificationtestingiscloselyrepresen-
sun, thermal radiative emission, and reflection from chamber
tative of that of the flight vehicle. The thermal balance test
walls (even at liquid nitrogen temperatures).
performed on this model gives the opportunity, once again, to
verify the thermal design and also to evaluate any changes 5.5.2 Prototype (Qualification) Test—The prototype space-
made due to thermal model test results. The test method here craftisnormallyusedforqualificationtests.Typicallyitisnear
includes exposure of the spacecraft to as realistic a space flightconfiguration,withallsubsystemscapableofperforming
environment as possible and also, perhaps, to some unrealistic their normal functions.The ideal qualification test will include
but readily definable thermal environments. The accurate some test exposures that are identical to those used on the
simulation of the space environment allows a determination of thermal control model. This provides a further verification of
in-spaceoperatingtemperatures.Thethermalinputsthatdonot the thermal design, particularly of any parts of the thermal
simulate space conditions may be used in some cases to subsystem modified as a result of thermal control model
determine the spacecraft thermal response. Perhaps the most testing. The most significant result of the qualification space-
important aspect of the qualification test is the verification of crafttestexposureisproofofthefunctionalperformanceofall
spacecraft functional operation while all components are at, or spacecraftsubsystems,inadditiontothethermalsubsystem.To
near, their in-space thermal conditions (both transient and achievethisend,andtodemonstratesystemdesignmargins,an
steady-state). environment is produced that thermally stresses all systems
more severely than they will be stressed by the anticipated
5.4.3 Flight Model (Acceptance Test)—The thermal balance
space conditions. In conjunction with the thermal stresses,
test on a flight spacecraft provides assurance of satisfactory
functionaldesignmarginsarealsoverifiedbyoperationathigh
operation in space. The purpose of the test is to indicate any
and low bus voltages and at various input signal threshold
deficiencies, either functional or thermal, that may only be
conditions.
recognizable under thermal-vacuum conditions. Frequently,
thistestisthefinalcheckofthethermalsystemsandspacecraft 5.5.3 Flight Model (Acceptance) Test—The final thermal
functional performance before launch. balance test is performed on flight spacecraft before launch.
The ideal test is one in which the simulated conditions are
5.5 The Ideal Thermal Balance Test Program—It is desir-
representative of all of those that will be experienced in flight.
able to outline a test program that will satisfy all test
Extremehot,cold,andtransientconditionsshouldbesimulated
objectives, and provide the highest possible confidence in the
as well as nominal operations. Again, the functional design
reliability of the spacecraft. This idealistic planning may be
margin, as represented by bus voltage and control signal
donewithoutconsideringmanyofthenormalrestraintssuchas
tolerances, is demonstrated concurrently with the verification
cost, schedule, and facility limitations. However, when the
of the thermal design. Ideally, this would be a long duration
restraints are imposed, the compromises, as discussed in 5.6,
test, and would include numerous temperature cycles from hot
tend to highlight those areas where deviations from this ideal
to cold extremes. This technique has a relatively high prob-
have been made. The method of implementation and the test
ability of exposing infant mortalities and marginal operations
results will be different for each model of the spacecraft, since
due to component parameter drift.
the test exposure is specifically arranged to satisfy the desired
objectives. 5.6 Tradeoff Considerations—It is not usually possible to
have as complete and rigorous a test program as the one
5.5.1 Thermal Control Model Test—The design of the ideal
described in 5.5.Among the restraints to be considered are the
thermal control model spacecraft test includes two test con-
costs, in terms of money and schedule, and, as detailed in
cepts. One of these test concepts involves the accurate simu-
Section7,thecharacteristicsandlimitationsoftheexistingtest
lation of all significant characteristics of the space
facilities,aswellasthenatureofthespacecraftanditsmission
environment, the orbital conditions, and the precise control of
parameters.
spacecraft operational modes. Since this concept leads to test
results that match the response that would be obtained under 5.6.1 Cost and Schedule—The cost per hour to operate a
real space flight conditions, an analytical (mathematical) ther- major environmental test facility must enter into each decision
mal model may not be necessary. A second test concept about the duration of test exposures. The more desirable long
involves a known deviation from accurate simulation of all durationtestsaremuchmorecostly.Costsincludenotonlythe
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environmental test facilities personnel and materials, but also characteristics, and spacecraft and mission parameters may be
the supporting spacecraft personnel and data reduction activi- prepared to assist in the final test definition. For a complete
ties. On flight spacecraft the space simulation test comes very systems integration test, this matrix is very complex and
lateintheintegrationsequence.Atthistimeinaspaceprogram certainly is beyond the scope of this recommended practice.
there is usually a considerable schedule urgency to meet a However, a matrix is provided in 7.6 for the thermal balance
launchdatecommitment.Thesecostandschedulefactorsmust testingphaseonly.Thefinaltestdefinitionisapyramidformed
be examined in terms of reliability as well as spacecraft by the many materials tests, subsystem tests, and supporting
requirements. For example, there are specific technical factors analysis which all provide confidence in meeting the overall
in addition to the subjective view that a longer test is a better objectives. Several examples of test facility configurations are
test. The thermal time constant of the spacecraft, that is, the given to illustrate special conditions which may influence the
time required to reach an equilibrium condition under a given test design.
set of thermal inputs, establishes a minimum duration for 5.7.1 Variable Solar Flux Vector—Most spacecraft do not
thermal design verification. For qualification and acceptance maintain a constant orientation with respect to the sun. The
spacecraft, this may be further extended by the minimum change in altitude may occur at the orbital period, seasonally,
length of time required to perform a complete spacecraft duringspacecraftmaneuvers,oratothertimesdependingupon
functional test. themissionprofile.Thesimulationofdifferentsolarfluxangles
may be accomplished by physically moving the spacecraft to
5.6.2 Facilities—The test facility itself provides the major
the desired position within the stationary solar beam. In some
influence on test tradeoffs and configuration. The size of the
instances, especially with spin-stabilized spacecraft, the me-
available chamber, the method of loading it (that is, top,
chanicalcomplexityofproducingavariable-spinaxishandling
bottom,side,andsoforth),andthedirectionofincidenceofthe
fixture precludes this approach. An equally effective test
solar simulator beam, are all important factors. Among other
methodusesamovablemirrortoredirectthesolarbeamtothe
things, these tend to determine the basic geometry of the
desired angle. Tests have been successfully performed in this
support fixture. The fixture design is also influenced by
manner using plane mirrors up to 100 ft in area. The use of a
spacecraft orbital characteristics such as spin rate and sun
remotelypositionablemirrorframemaypermitthestimulation
angles, and by thermal influences, including conduction errors
of summer, equinox, and winter incident angles on a spinning,
into and out of the fixturing and shadowing from various
geosynchronous spacecraft without returning the chamber to
sources.Thesolar-simulatorcharacteristicsmustbethoroughly
atmospheric pressure.
understood to allow proper test evaluation. Major factors are
5.7.2 Stationary Test of Spinning Spacecraft—It is some-
spectrum,total-beamirradiance,uniformityofirradianceinthe
timesnecessarytoperformastationarytestonaspacecraftthat
total test volume, solar beam divergence angle, and temporal
is designed to spin in orbit. An example of this is a commu-
variations. These factors, together with recommended
nicationssatelliteonwhichthetranspondersmustbeconnected
tradeoffs, are discussed in 7.2 and 7.5.
to the test equipment by waveguides or coaxial cables, which
5.6.3 Spacecraft and Mission Parameters—Each spacecraft
precludestheuseofsliprings.Thisthermalbalancetestmaybe
andeachmissionpresentsuniquecharacteristicswhichmustbe
accomplished by a circumferential tungsten lamp array.
considered in the design of the test exposure. For attitude-
5.7.3 Combined Solar Sources—Acombination of tungsten
stabilized planet-oribiting spacecraft, the orientation with re-
or infrared sources may have to be used in conjunction with a
specttorunandplanethasconsiderablethermalinfluence.The
spectrally accurate source if the high quality source does not
altitude of the orbit determines the amount of albedo and earth
irradiate a large enough area.Whenever this technique is used,
emissionthatmustbesimulatedoraccountedfor.Thestructure
it is essential to consider all of the effects of the differences
of the spacecraft also has an effect in the amount of self-
between the sources in spectrum, subtense angle, and diver-
shadowing by appendages and solar paddles. Along this same
gence angle. These aspects are discussed more thoroughly in
line there may be extraneous heat sources. An example is the
7.1 and 8.5.1.
use of nuclear generators for power sources on deep-space
missions.Therearesomespacecraft,orspacecraftsubsystems,
6. Safety Considerations
in which the test item surface temperature is so high (for
6.1 Purpose—The purpose of this section is to recommend
example several hundred degrees Fahrenheit) that it may be
procedures that will help to ensure the safety of persons
necessary to use a liquid nitrogen temperature cold wall in the
(includingcasualobservers)associatedwiththeuse,operation,
chamber. All of these things are considered in the tradeoffs
and maintenance of solar simulators.
which lead to a optimum test design. 7.3 and 7.4 cover the
subject in more detail.
6.2 Scope—Potentialhazardsarediscussedintermsofwhat
they are, their damage or consequences, and their exposure
5.7 Final Test Definition—The final test plan should be
rates and times (where applicable). The hazards have been
evaluated in terms of test adequacy after careful consideration
categorized into mechanical, chemical, electrical, radiation,
of the objectives and facility capabilities. In the case of the
thermal,andmiscellaneoushazards.Thepreventionofhazards
thermalcontrolmodeltest,theevaluationconsistsofassessing
and the protection and care of the victims are also discussed.
the fidelity of the space simulation and the completeness and
Onlythosehazardsandinjuriespeculiartosolarsimulationare
accuracy of the instrumentation. The qualification and accep-
included.
tance tests pose a somewhat more complex problem since all
subsystems must be tested.Amatrix of test objectives, facility 6.3 General Instructions:
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6.3.1 Whenever a solar simulator, laser, or similar equip- 6.6 Discussion of Hazards:
ment is being operated, suitable warning signs should be
6.6.1 Mechanical Hazards—Mechanical hazards involve
clearly displayed at all entrances to the work area.Acomplete
thosehazardswhichcouldproducephysicalinjurytopersonnel
list of safety procedures appropriate to the facility should be
or equipment. They can be caused by exploding high-pressure
clearly and prominently displayed.
lamps, implosion of vacuum windows, falls, ruptures, high-
6.3.2 Every person who may be operating in the work area
pressure systems, structural hazards, lifting and handling,
should be informed to the hazards involved, safety precautions
rotating machinery, and so forth.
to be taken, and supervisory or medical personnel to be
6.6.1.1 Exploding High-Pressure Lamps—A compact arc
contactedincaseofaccidents.Alloperationalpersonnelshould
lamp, when in use, is at a high internal pressure
...