Operational phase of the inflatable radiator; inflatable space structures; advantages; inflated thin-film satellites; antenna configuration; 3 meter diameter test paraboloid (HAIR program); and weight breakdown for the 100 meter diameter reflector are outlined. This presentation is represented by viewgraphs only

Williams, Geoff

English

NASA Technical Reports Server

N90-21799
INFLATABLE ANTENNAS FOR MICROWAVE POWER TRANSMISSION
Geoff Williams
L'Garde Inc.
Tustin, California 92680
L'GARDE SPACE INFLATABLES
Various ABRES (IEO, DDT, AMT, AREP, TREP)
(1971 through 1978)
Sounding rocket and
ICBM flights
Space Defense
ITV (SD)
Have Busk I (RPL)
Study
Study
Reflectors
Antenna (NASA)
Antenna (Martln)
VOA (TRW/State)
Solar Concentrator (RPL)
Study
Study
Study
Ground test
SDI - related
Radiator (AFWAL SDIO)
NP. Beam (SAIC LMSC)
SRMP (BMO/USASDC)
STEP SD (In evaluation)
Ground test
Flight test
Flight test
Flight test
PRECEDING PAGE BLANK NOT FILMED
51
https://ntrs.nasa.gov/search.jsp?R=19900012483 2020-03-19T23:23:18+00:00Z
INFLATABLE SPACE STRUCTURES
NASA - antennas
USAF-BMO - decoys
USAF-SD - ITV
AFRPL - Solar collector
AFWAL - radiator
SDIO - radiator
IJSASDC - decoys
Aerospace Companies: (TRW, Martin,
MDAC, Avco, MIT/LL)
L37
52
OPERATIONAL PHASE OF THE INFLATABLE RADIATOR
This picture shows the operations phases of a weapon system using the High Power Inflatable
Radiator Systems (HIRS) and the insert shows the design of an inflatable radiator. The operation of
HIRS is described as follows: In its normal standby phase the inflatable radiator is housed inside
the space weapon system to enhance its survivability and to keep it at a temperature where the
liquid water, used in the system, will not freeze. In the alert phase a small amount of water vapor
is used to inflate the radiator fabric bag at low pressure. This is done to keep the bag from
bursting as a result of the high water vapor pressure impulse developed in a short time during the
firing phase. In the firing phase the radiator is designed to operate for 2 minutes continuously.
This is based on a polar orbit of 100 minutes where the radiator will be ready for use 4 minutes of
this time. The radiator system will be capable of dissipation between 10 and 100 megawatts (mw)
levels of waste heat. During the cooling phase the waste heat will be dissipated in space by
condensation on the inside of the fabric bag. As the vapor pressure subsides a sensor will actuate
the retraction mechanism to bring in the bag. A 5.6 psi pressure will be maintained in the bag
during retraction in order to keep the bag stable, to facilitate folding the fabric radiator, and to
prevent freezing of water on the cold side of the bag. After all the heat is dissipated the bag
will be folded and stored in the temperature controlled environment of the spacecraft. The water
vapor will have condensed, and the liquid will have been collected and returned to a reservoir to
use again.
In the insert the inflatable radiator is shown partially retracted. The fabric bag will be
made of a material which has low water permeability, a fabric and seam strength of 500 Ibs/linear
in., high emissitivity and a low contamination level factor for the space surrounding the weapon
system. Water recovery is done by using a sponge to gather the water which has condensed on the
sides of the radiator, and then squeezing the sponge into a recovery system which transmits it to a
reservoir to be used again. The friction rollers and motor are used to roll in the fabric radiator
and then fold it into the spacecraft.
53
INTERNAL VIEW OF THE INFLATABLE RADIATOR DURING WEAPON FIRING PHASE
This picture shows the internal view of the inflatable radiator during the space weapon's firing phase. The waste heat from
the weapon is absorbed by the boiler tubes surrounding it. The water is turned into water vapor at 7S°C and then is forced
through a vapor/liquid separator where the vapor is directed toward the inflatable fabric radiator and the water is returned to
the reservoir. The vapor pressurizes the radiator, which has previously been inflated at low pressure to avoid destroying it
during the actual firing phase. The vapor condenses on the radiator wall where it is then sponged up, recovered and returned to
the reservoir to be used again. The operational characteristics fDr the inflatable radiator are as follows.
HIRS TECHNICAL REQUIREMENTS
l° OPERATIONAL RADIATOR DIMENSIONS
o Die - 8'
o Helght - 40'
o Area input opening to radiator is I/4 total radiator cross sectional area
o Time expanded in polar orbit Is I00 minutes
o Operational time 4 mln
o Continuous operations 2 min
2. FABRIC REQUIREMENTS
o Water permeability - as low as possible
o Tensile strength - 500 Ibs/llnear inch
o Temperature - 75°C
o G Force - max of .IG in space
o Launch forces will be same as the space shuttle
o Withstand radiation intensity from the sun as a 6000K Black Body at 1350 W/M 2
o Minimum contamination leakage
o Internal pressure - Max 5.6 psi
o Compatible with water vapor and liquid for long periods of time (I0 yr.)
RADIATOR must be capable of expanding and contracting in space a minimum of 25 times.
o Test once a year for lO years
o Potential use as a weapon is IS times
4. SYSTEM WASTE ,HEAT REQUIREMENTS
o Operation - I0 to TO0 MW
S. TRANSPORTATION - Must be collapsable as cargo to fit into the space shuttle.
B. WATER RESERVOIR/BAG - IOKG
7. HEAT RATE DISSIPATION TO SPACE 3.8 KILOWATTS
54
ORIGINAL PAGE
BLACK AND WHITE PHOTOGRAPH
, ,/? J ,
Of POOR r_/,ttTY
INFLATABLE SPACE TARGET
SOLAR THERHAL ROCKET
C 89 12822
55
SOLAR THERNAL ROCKET - L'GARDE DESIGN
INFLATED REFLECTOR CONCEPT
56
NASA INFLATED THIN-FILM SATELLITES
SYSTEM
EXPLORER IX
EXPLORER Xl_
ECHO II
PAGEOS I
PWT
(LB)
135
34
34
58O
149
DIA
(FT)
100
12
12
135
100
LAUNCH ILIFEI PURPOSE
DATE (YEARS)
I
AUG 1960
FEB 1961
DEC 1963
JAN 1964
JUNE 1966
COMM.
HI-ALT.
DENSITY
HI-ALT.
DENSITY
COMM.
EARTH
SURVEY
FILM/METAL
THICKNESS (MIL)
0.5 FILM/VDA
.5 FILM (2)
,5 AL (2)
.5 FII_M (2)
.5 AL (2:'
.35 FILM
.16 AL (2)
.5 FILM
VDA (2000 /_)
ADVANTAGES OF INFLATABLE REFLECTOR
CONFIGURATION •SUPPORTED AT POWER UNIT
• RADIATOR CLOSE TO POWER UNIT
DYNAMICS •MASS CONCENTRATED AT GIMBAL
• GOOD DAMPING OF MOTION
IGENERALLY I_!OT SIMPLE HARMONIC MOTION
RELIABILITY •TOTAL INFLAT/kBLE PACKAGE EASILY REPLACED
• INFLATABLES HAVE SPACE EXPERIENCE
iGROUND TESTABLE
• LOW WEIGHT AND VOLUME
l LOW ENGINEERING/MANUFACTURING COST
• DEVELOPMENT ALREADY STARTED
• EVA MINIMIZED
• INCREASING INTERCEPT AREA IS MINOR CHANGE
I TUNINQ POSSIBLE THRU PRESSURE ADJSTMT.
PRACTICALITY
PERFORMANCE
57
ANTENNA CONFIGURATION
FEED LOCATION
/--'INFLATED CONE
INFLATED ___D
O.006mm VDA FILM IZING TORUS
0.06 mm COMPOSITE
NEAR TERM APPLICATIONS FOR
LARGE INFLATABLE ANTENNAS
U, S. STATE DEPARTMENT
SYSTEM
VOICE OF AMERICA
COMMENTS
700 FT. DIA., 26 MHz
NASA GROUND MOBILE 200 FT. DIA., 850 MHZ
U, S, NAVY INTEGRATED TACTICAL
SURVEILLANCE SYSTEM
SYNCHRONOUS VERSION, 200 FT, _;A,
U, S, AIR FORCE SPACE RADAR
MM-WAVE RADIOMETER
SOLAR CONCENTRATOR
"DARK" PROGRAMS
PHASED ARRAYS PRIME CANDIDATE
10-30M, BETTER ACCURACY NEEDED
SOLAR ROCKET
50-250 FT. DIA,
58
3 - METER INFLATED REFLECTOR - UNINFLATED
CONDITION,DEVELOPED FOR NASA LANGLEY
3 - METER INFLATED REFLECTOR -
INFLATED CONDITION
C-89-12826
ORIGINAL p _,_. ..
Of POOR Q_L;ALITY
59
GOREDESIGN
ON-AXISREFLECTOR
.,,,.--
DETERMINE SHAPES OF THE GORES THAT WILL
FORM A PARABOLIC DISH UPON INFLATION
"FATTER" THAN A TRIANGULAR WEDGE
"* USE FLATE CODE "*
O.NE M ONE METER DIAMETER INFLATED REFLECTOR DEVELOPED FOR THE B
AF ROCKET PROPULSION LAB
60
ORIGrNAL PAGE
BLACK AND WHITE PHOTOGRAPH
3-Meter-Diameter Test Paraboloid
(HAIR Program)
!
ORIGINAL PAGF.. 13
OF POOR QUALITY
61
_RrG!NAI2 Pa6E
BLACK AND WHITE PHOi-OGRAPH
0PT ICAL
LASER
TOP VIEW r L
VACUUM CHP/4BER HOUSING
XR
2.05 MM
L3.83 MM
'-- S i i
MOVABLE SCALE
LASER
YR
SIDE VIEW
SLOPE-MEASUREMENT OPTICAL SETUP
62
A|
E
,Y
O
m,
q2
U
e"
e•
i
N
-50
-I00
-150
MEASUREMENT OF PARABOLOID SURFACE ACCURACY
32 i-mil VDA Mylar Gores
½-mil Mylar Tape
3 Meter Diameter
Measured 673mm Below Centerline
m
500
i
-I000
X2 + y2 . (D/2)2
Z- _f
. X2 + (673)2 - (15oo)2
4(2963)
I
; /
-500 O 500
i
1000 1500
X - Horizonta| Distance (mm)
2 w
MEASUREMENT OF PARABOLOID
v
4_
L¢_
&.
%.
W-
C
O
e•
(D
i
0 •
-i
-Z l i
-1500
-1000
ee e
Q•
t
0.76 mm RMS
• • •e
• •e •
i
t
-500 O 500
X- Horizontal distance (,m)
1000 1500
63
0t.
0
i0,000
1,000
100
10
-I
.01
REQUIRED ANTENNA ACCURACY
\
\
\
\ \
\\ \
XII o -
required surface accuracy
\
\
\
- Wavelength,X
\
\
\
\\
\
\
\
\
\\
\
\
I "\
I
' ' I ' i
Mhz 10 100 1 GH z 10 100
Microwave
I
I I
,,, \
1 Thz 10
Frequency
CYLINDER TESTS
THESE RIGIDIZED CYLINDERS WILL FORM THE TORUS OF THE INFLATED ANTENNA.
BEFOREPRESSURE
ORIGINAL PAGE IS
OF POORQUALITY
AFTER PSIG
64
AFTER 49.5 LBS FORCE COMPRESSIVE
ORIGINAL PAGE
BLACK AND WHITE PHOTOGRAPH
WEIGHT BREAKDOWN FOR 100-METER
DIAMETER REFLECTOR
_E
O
,,i
I..-
-r-
]000-
300-
100-
30-
10-
3-
1
O
TOTAL
TORUS
......
"'"-WATER FOR PRESSURE MAINTENANCE
WATER FOR YIELDING PARABOLOID
INFLATION SYSTEM HARDWARE
CARBON DIOXIDE
I 1 I
0.5 I.O ].5
FOCAL LENGTH/DIAMETER
I
2.0
I •
PROGRAM PLAN
1.0 mm accuracy
- have demonstrated this for NASA and AFAL
- need to incorporate torus
- 1.5 years to build I0 meter flight reflector
- roughly $2 to $3 million
1 0.1 mm accuracy
- best tests have yielded 0.i mm
need higher pressures, heavier torus
- 3 meter testing
- three years to build i0 meter flight reflector
- roughly $5 million
65



Inflatable antennas for microwave pwoer transmission

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