ASTM E491-73(1999)
(Practice)Standard Practice for Solar Simulation for Thermal Balance Testing of Spacecraft
Standard Practice for Solar Simulation for Thermal Balance Testing of Spacecraft
SCOPE
1.1 Purpose:
1.1.1 The primary purpose of this practice is to provide guidance for making adequate thermal balance tests of spacecraft and components where solar simulation has been determined to be the applicable method. Careful adherence to this document should ensure the adequate simulation of the radiation environment of space for thermal tests of space vehicles.
1.1.2 A corollary purpose is to provide the proper test environment for systems-integration tests of space vehicles. An accurate space-simulation test for thermal balance generally will provide a good environment for operating all electrical and mechanical systems in their various mission modes to determine interferences within the complete system. Although adherence to this practice will provide the correct thermal environment for this type of test, there is no discussion of the extensive electronic equipment and procedures required to support systems-integration testing.
1.2 Nonapplicability -This practice does not apply to or provide incomplete coverage of the following types of tests:
1.2.1 Launch phase or atmospheric reentry of space vehicles,
1.2.2 Landers on planet surfaces,
1.2.3 Degradation of thermal coatings,
1.2.4 Increased friction in space of mechanical devices, sometimes called "cold welding,"
1.2.5 Sun sensors,
1.2.6 Man in space,
1.2.7 Energy conversion devices, and
1.2.8 Tests of components for leaks, outgassing, radiation damage, or bulk thermal properties.
1.3 Range of Application:
1.3.1 The extreme diversification of space-craft, design philosophies, and analytical effort makes the preparation of a brief, concise document impossible. Because of this, various spacecraft parameters are classified and related to the important characteristic of space simulators in a chart in 7.6.
1.3.2 The ultimate result of the thermal balance test is to prove the thermal design to the satisfaction of the thermal designers. Flexibility must be provided to them to trade off additional analytical effort for simulator short-comings. The combination of a comprehensive thermal-analytical model, modern computers, and a competent team of analysts greatly reduces the requirements for accuracy of space simulation.
1.4 Utility -This recommended practice will be useful during space vehicle test phases from the development through flight acceptance test. It should provide guidance for space simulation testing early in the design phase of thermal control models of subsystems and spacecraft. Flight spacecraft frequently are tested before launch. Occasionally, tests are made in a space chamber after a sister spacecraft is launched as an aid in analyzing anomalies that occur in space.
1.5 This standard does not purport to address all of the safety problems, if any, associated with its use. It is the responsibility of the user of this standard to establish appropriate safety and health practices and determine the applicability of regulatory limitations prior to use.
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Standards Content (Sample)
NOTICE: This standard has either been superseded and replaced by a new version or withdrawn.
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Designation: E 491 – 73 (Reapproved 1999)
Standard Practice for
Solar Simulation for Thermal Balance Testing of Spacecraft
This standard is issued under the fixed designation E 491; the number immediately following the designation indicates the year of
original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval. A
superscript epsilon (e) indicates an editorial change since the last revision or reapproval.
1. Scope designers. Flexibility must be provided to them to trade off
additional analytical effort for simulator shortcomings. The
1.1 Purpose:
combination of a comprehensive thermal-analytical model,
1.1.1 The primary purpose of this practice is to provide
modern computers, and a competent team of analysts greatly
guidance for making adequate thermal balance tests of space-
reduces the requirements for accuracy of space simulation.
craft and components where solar simulation has been deter-
1.4 Utility—This recommended practice will be useful dur-
mined to be the applicable method. Careful adherence to this
ing space vehicle test phases from the development through
document should ensure the adequate simulation of the radia-
flight acceptance test. It should provide guidance for space
tion environment of space for thermal tests of space vehicles.
simulation testing early in the design phase of thermal control
1.1.2 A corollary purpose is to provide the proper test
models of subsystems and spacecraft. Flight spacecraft fre-
environment for systems-integration tests of space vehicles. An
quently are tested before launch. Occasionally, tests are made
accurate space-simulation test for thermal balance generally
in a space chamber after a sister spacecraft is launched as an
will provide a good environment for operating all electrical and
aid in analyzing anomalies that occur in space.
mechanical systems in their various mission modes to deter-
1.5 This standard does not purport to address all of the
mine interferences within the complete system. Although
safety concerns, if any, associated with its use. It is the
adherence to this practice will provide the correct thermal
responsibility of the user of this standard to establish appro-
environment for this type of test, there is no discussion of the
priate safety and health practices and determine the applica-
extensive electronic equipment and procedures required to
bility of regulatory limitations prior to use.
support systems-integration testing.
1.2 Nonapplicability—This practice does not apply to or
2. Referenced Documents
provide incomplete coverage of the following types of tests:
2.1 ASTM Standards:
1.2.1 Launch phase or atmospheric reentry of space ve-
E 259 Practice for Preparation of Pressed Powder White
hicles,
Reflectance Factor Transfer Standards for Hemispherical
1.2.2 Landers on planet surfaces,
and Bi-Directional Geometries
1.2.3 Degradation of thermal coatings,
E 296 Practices for Ionization Gage Application to Space
1.2.4 Increased friction in space of mechanical devices,
Simulators
sometimes called “cold welding,”
E 297 Methods for Calibrating Ionization Vacuum Gage
1.2.5 Sun sensors,
Tubes
1.2.6 Man in space,
E 349 Terminology Relating to Space Simulation
1.2.7 Energy conversion devices, and
2.2 ISO Standard:
1.2.8 Tests of components for leaks, outgassing, radiation
ISO 1000-1973 SI Units and Recommendations for the Use
damage, or bulk thermal properties.
of Their Multiples and of Certain Other Units
1.3 Range of Application:
2.3 American National Standards:
1.3.1 The extreme diversification of space-craft, design
ANSI Y10.18-1967 Letter Symbols for Illuminating Engi-
philosophies, and analytical effort makes the preparation of a
neering
brief, concise document impossible. Because of this, various
ANSI Z7.1-1967 Standard Nomenclature and Definitions
spacecraft parameters are classified and related to the important
for Illuminating Engineering
characteristic of space simulators in a chart in 7.6.
ANSI Y10.19-1969 Letter Symbols for Units Used in Sci-
1.3.2 The ultimate result of the thermal balance test is to
ence and Technology
prove the thermal design to the satisfaction of the thermal
Annual Book of ASTM Standards, Vol 06.01.
1 3
This practice is under the jurisdiction of ASTM Committee E-21 on Space Annual Book of ASTM Standards, Vol 15.03.
Simulation and Applications of Space Technology and is the direct responsibility of Discontinued, see 1984 Annual Book of ASTM Standards, Vol 15.03.
Subcommittee E21.04 on Space Simulation Test Methods. Available from the American National Standards Institute, 11 W. 42nd St., 13th
Current edition approved Sept. 27, 1973. Published November 1973. Floor, New York, NY 10036.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959, United States.
E 491
3. Terminology 3.2.10 collimation angle—in solar simulation, the angular
nonparallelism of the solar beam, that is, the decollimation
3.1 Definitions, Symbols, Units, and Constants—This sec-
angle. In general, a collimated solar simulator uses an optical
tion contains the recommended definitions, symbols, units, and
component to image at infinity an apparent source (pseudo sun)
constants for use in solar simulation for thermal balance testing
of finite size. The angle subtended by the apparent source to the
of spacecraft. The International System of Units (SI) and
final optical component referred to as the collimator, is defined
International and American National Standards have been
as the solar subtense angle and establishes the nominal angle of
adhered to as much as possible. Definitions E 349 is also used
decollimation. A primary property of the “collimated” system
and is so indicated in the text. Table 1 provides commonly used
is the near constancy of the angular subtense angle as viewed
symbols.
from any point in the test volume. The solar subtense angle is
3.2 Definitions:
therefore a measure of the nonparallelism of the beam. To
3.2.1 absorptance (a , a ,a )—ratio of the absorbed radiant
e v
avoid confusion between various scientific fields, the use of
or luminous flux to the incident flux (E 349) (Table 1).
solar subtense angle instead of collimation angle or decollima-
3.2.2 absorptivity of an absorbing material—internal ab-
tion angle is encouraged (see solar subtense angle).
sorptance of a layer of the material such that the path of the
3.2.11 collimator—an optical device which renders rays of
radiation is of unit length (E 349).
light parallel.
3.2.3 air mass one (AM1)—the equivalent atmospheric
3.2.12 decollimation angle—not recommended (see colli-
attenuation of the electromagnetic spectrum to modify the solar
mation angle).
irradiance as measured at one astronomical unit from the sum
3.2.13 diffuse reflector—a body that reflects radiant energy
outside the sensible atmosphere to that received at sea level,
in such a manner that the reflected energy may be treated as if
when the sun is in the zenith position.
it were being emitted (radiated) in accordance with Lambert’s
3.2.4 air mass zero (AM0)—the absence of atmospheric
law. The radiant intensity reflected in any direction from a unit
attenuation of the solar irradiance at one astronomical unit
area of such a reflector varies as the cosine of the angle
from the sun.
between the normal to the surface and the direction of the
3.2.5 albedo—the ratio of the amount of electromagnetic
reflected radiant energy (E 349).
radiation reflected by a body to the amount incident upon it.
3.2.14 dispersion function (X/l)—a measure of the separa-
3.2.6 apparent source—the minimum area of the final
tion of wavelengths from each other at the exit slit of the
elements of the solar optical system from which issues 95 % or
monochromator, where X is the distance in the slit plane and l
more of the energy that strikes an arbitrary point on the test
is wavelength. The dispersion function is, in general, different
specimen.
for each monochromator design and is usually available from
3.2.7 astronomical unit (AU)—a unit of length defined as
the manufacturer.
the mean distance from the earth to the sun (that is,
3.2.15 divergence angle—see solar beam divergence angle
149 597 890 6 500 km).
(3.2.60).
3.2.8 blackbody (USA), Planckian radiator—a thermal ra-
diator which completely absorbs all incident radiation, what- 3.2.16 electromagnetic spectrum—the ordered array of
ever the wavelength, the direction of incidence, or the polar- known electromagnetic radiations, extending from the shortest
ization. This radiator has, for any wavelength, the maximum wavelengths, gamma rays, through X rays, ultraviolet radia-
spectral concentration of radiant exitance at a given tempera- tion, visible radiation, infrared and including microwave and
all other wavelengths of radio energy (E 349).
ture (E 349).
3.2.9 collimate—to render parallel, (for example, rays of 3.2.17 emissivity of a thermal radiator e, e5 M /M
e,th e
light). (e5 1)—ratio of the thermal radiant exitance of the radiator to
TABLE 1 Commonly Used Symbols
Symbol Quantity Definition Equation or Value Unit Unit Symbol
Q radiant energy, work, joule J
quantity of heat
−1
F radiant flux F5 dQ/dt watt (joule/second) W, Js
−2
E irradiance (receiver) flux E 5 dF/dA watt per square metre W·m
density
−2
M radiant exitance (source) M 5 dF/dA watt per square metre W·m
−1
I radiant intensity (source) I 5 dF/dv watt per steradian W·sr
v5 solid angle through which flux from source is radiated
−1 −2
L radiance L 5 dI/(dA cosu ) watt per steradian 5 W·sr ·m
square metre
u5 angle between line of sight and normal to surface dA
t transmittance t5F, transmitted/F, incident none
t(l) spectral transmittance t(l) 5F(l), transmitted/F(l), incident none
r reflectance (total) r5F, reflected/F, incident none
eH emittance (total eH 5 M, specimen/M, blackbody
hemispherical)
a absorptance a5F, absorbed/F, incident none
a solar absorptance a 5 solar irradiance absorbed/solar irradiance incident none
s s
E 491
that of a full radiator at the same temperature, formerly 3.2.31 irradiance, spectral [E or E(l)]—the irradiance at a
l
“pouvoir emissif ” (E 349). specific wavelength over a narrow bandwidth, or as a function
of wavelength.
3.2.18 emittance (e)—the ratio of the radiant exitance of a
specimen to that emitted by a blackbody radiator at the same
3.2.32 irradiance, temporal—the temporal variation of in-
temperature identically viewed. The term generally refers to a
dividual irradiances from the mean irradiance. The temporal
specific sample or measurement of a specific sample. Total
variations should be measured over time intervals equal to the
hemispherical emittance is the energy emitted over the hemi-
thermal time constants of the components. The temporal
sphere above emitting element for all wavelengths. Normal
stability of total irradiance can be defined as:
emittance refers to the emittance normal to the surface to the
¯
E 56100@~DE 1DE !/2E# (2)
emitting body. t t ~min! t ~max!
3.2.19 exitance at a point on a surface (radiant exitance)
3.2.33 irradiance, total—the integration over all wave-
(M)—quotient of the radiant flux leaving an element of the
lengths of the spectral irradiance.
surface containing the point, by the area of that element,
3.2.34 irradiance, uniformity of—uniformity of total irradi-
−2
measured in W·m (E 349) (Table 1).
ance can be defined as:
3.2.20 field angle—not recommended (see solar beam sub-
¯
E 56100@~E 1 E !/2E# (3)
tense angle). u ~min! ~max!
3.2.21 flight model—an operational flight-capable space-
where:
craft that is usually subjected to acceptance tests.
E 5 uniformity of the irradiance within the test vol-
u
3.2.22 flux (radiant, particulate, and so forth)—for electro-
ume, expressed as a percent of the mean irradi-
magnetic radiation, the quantity of radiant energy flowing per
ance,
unit time; for particles and photons, the number of particles or
E 5 smallest value obtained for irradiance within the
(min)
photons flowing per unit time (E 349).
test volume, and
E 5 largest value obtained for irradiance within the
3.2.23 gray body—a body for which the spectral emittance
(max)
test volume.
and absorptance is constant and independent of wavelength.
The term is also used to describe bodies whose spectral
Uniformity of irradiance values must always be specified
emittance and absorptance are constant within a given wave-
together with the largest linear dimension of the detector used.
length band of interest (E 349).
3.2.35 Lambert’s law—the radiant intensity (flux per unit
3.2.24 incident angle—the angle at which a ray of energy
solid angle) emitted in any direction from a unit-radiating
impinges upon a surface, usually measured between the direc-
surface varies as the cosine of the angle between the normal to
tion of propagation of the energy and a perpendicular to the
the surface and the direction of the radiation (also called
surface at the point of impingement or incidence.
Lambert’s cosine law). Lambert’s law is not obeyed exactly by
3.2.25 infrared radiation—see electromagnetic spectrum
most real surfaces, but an ideal blackbody emits according to
(E 349).
this law. This law is also satisfied (by definition) by the
3.2.26 insolation—direct solar irradiance received at a sur-
distribution of radiation from a perfectly diffuse radiator and by
face, contracted from incoming solar radiation.
the radiation reflected by a perfectly diffuse reflector. In
3.2.27 integrating (Ulbrecht) sphere—part of an integrating
accordance with Lambert’s law, an incandescent spherical
photometer. It is a sphere which is coated internally with a
blackbody when viewed from a distance appears to be a
white diffusing paint as nonselective as possible, and which is
uniformly illuminated disk. This law does not take into account
provided with associated equipment for making a photometric
any effects that may alter the radiation after it leaves the
measurement at a point of the inner surface of the sphere. A
source.
screen placed inside the sphere prevents the point under
3.2.36 maximum test plane divergence angle—the angle
observation from receiving any radiation directly from the
between the extreme ray from the apparent source and the test
source (E 349).
plane. This applies principally to direct projection beams
3.2.28 intensity—see radiant intensity.
where it is equivalent to one half the projection cone angle (see
3.2.29 irradiance at a point on a surface E ,E;E 5 dF /
e e e
Fig. 1).
dA—quotient of the radiant flux incident on an element of the
3.2.37 natural bandwidth—the width at half height of a
surface containing the point, by the area of that element
radiation source emission peak. It is independent of instrument
−2
me
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