Carbon phenolic heat shield technology is reviewed. Heat shield results from the outer planetary probe mission studies are summarized along with results of plasma jet testing of carbon phenolic conducted in a ten megawatt facility. Missile flight data is applied to planetary entry conditions. A carbon phenolic heat shield material is utilized and tailored to accommodate each of the probe missions. An integral heat shield approach is selected over in order to eliminate a high temperature interface problem and permit direct bonding of the carbon phenolic to the structural honeycomb sandwich. The sandwich is filled with a very fine powder to minimize degradation of its insulation properties by the high conductive hydrogen/helium gases during the long atmospheric descent phase

Mezines, S.

English

NASA Technical Reports Server

_" i_ _ _ i
•i
CARBON PHENOLIC HEAT SHIELDS FOR JUPITER/SATURN/URANUS
ENTRY PROBES
S. Mezines
McDonnell-Douglas Astronautics Company
MR. MEZINES: I am going to limit my talk to carbon pheno-
lic heat shield technology. As you probably know, these materials
have been around for a number of years and we have assimilated a
lot of fabrication and flight experience on these materials from
our numerous RV programs.
In this presentation I am going to cover three areas. First
of all, I will summarize the heat shield results from the outer
planetary probe mission studies that we've done in the last couple
of years. Secondly, I will attempt to demonstrate the applica-
bility of missile flight data to planetary entry conditions; and
finally, I will summarize the results of some recent plasma jet
testing of carbon phenolic conducted in our ten megawatt facility.
. .(!
: j( :_'
t
Figure 6-1 illustrates the common probe design that we have
developed for exploration of the outer planets. We propose to
use a carbon phenolic heat shield material and tailor the thick-
ness of the material to accommodate each of the probe missions.
We have selected an integral heat shield approach over
concepts utilizing an intermediate insulation layer in order to
eliminate a high temperature interface problem and permit direct
bonding of the carbon phenolic to the structural honeycomb sand-
wich. The sandwich is filled with a very fine powder to minimize
degradation of its insulation properties by the high conductive
hydrogen/helium gases during the long atmospheric descent phase.
The inner portion of the forebody heat shield has been hollowed
out to reduce both weight and heat conduction.
The afterbody heat shield is made of a low density elasto-
meric material which is light-weight and RF transparent.
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https://ntrs.nasa.gov/search.jsp?R=19750012314 2020-03-19T21:09:45+00:00Z
Forebody Heat Shield Afterbody Heat Shield
POLYURETHANE "
FOAM AFT COVER _0.25 IN.
PEAKS.
0OF
2500°F
HONEYCOMB fp= 18LB/FT3_
(POWDER FILLED_
HOLLOWED CARBON PHENOLIC iF,=36LB/FT3_
(FIBERFILLED)
CARBON (p= 90LB/FT3_
PHENOLIC
_0.25 IN.
IRF TRANSPARENT1
.:.2:, ::" .';
2 -
t _ : ,'
•" " -,x_.:z J
c T:.: .::_C: ,/:,'{
Figure 6-1. Planetary Probe Heatshield
Figure 6-2 depicts the convective and radiative heat flux
associated with entry into each of the planets. As indicated,
the fluxes are very high, in the 40,000 to 50,000 BTU/FT 2 sec
range, and predominantly radiative. These fluxes and the heat
shield requirements to be shown later were computed by Aerotherm
Corporation under contract to NASA Ames.
The magnitude of heating associated with each planetary
entry is very strongly influenced by the initial entry angle and
atmospheric mode/assumed. For instance, steep entries into the
cold atmospheres of Saturn and Uranus result in heating rates as
high as those encountered in a shallow entry into the Jupiter
nominal atmosphere, even though the entry velocity at Jupiter is
50 percent higher than entry into the other planets.
VI-3
40- •
'- !:
; _ QT =336,500
• ::..__i ,..,_ 30- 00__/
_:_ "' .n°,.=x20| iR = 190,
TIME FROMENTRY, t E - sec
• NO BLOWING
• STAGNATIONPOINT
STEEP - URANUS
COOLATMOSPHERE
YI = -50°
VR = 26.3 krn/sec •
,-QT=95,000
/--QR =64,000
• 5o
40 ' _I 40
30; 30
::
I0
o o
16 20 24
TIME FROMENTRY- sec
STEEP- SATURN
COOLATMOSPHERE
YI = -400
VR = 30.4 km/sec
QT=53,500
16 20 24
TIME FROM ENTRY -sec
Figure 6-2. Planetary Entry Heating Environments
= . ,,:r
C
. - .,;
.... • &
m
The high heating rate for the Uranus entry is due to the
large proportion of helium dictated by the cold atmospheric
model. The high helium/hydrogen ratio results in not only a
higher deceleration load and stagnation pressure but also in
higher shock layer temperatures and much higher radiation fluxes.
Selection of the shallow Jupiter entry condition was made on the
basis of the preliminary Pioneer i0 data which indicated that the
atmosphere composition is near the solar abundance ratio (nom-
inal model) and better knowledge of the planet's ephemeris data
permit shallow entry with very small uncertainty in entry angle.
Heat shield thickness requirements for each of the outer
planets is established by analyzing a number of critical entry
VI-4
iii
i'•_;ii
... -.:
trajectories which bound the entry envelope and atmospheric model
uncertainty. In general, steep entries coupled with the cold at-
mospheres model definition results in high heating rates and high
surface recession rates whereas shallow entries and warm atmosphere
lead to milder heating rates but longer durations and higher insul-
ation requirements.
For Saturn, the shallow-warm atmosphere entry sized the heat
shield even though the peak heat flux was only 2300 BTU/FT2-sec
and practically no surface recession occurred. Conversely, entry
into the Uranus cold-dense atmosphere model results in very high
heating rates so that material recession sizes the heat shield
thickness requirements. For Jupiter, we have purposely limited
the entry angle to very shallow values (about 7.5 °) in order to
alleviate the heating and heat shield requirements. Furthermore,
the Pioneer 10 data indicate an atmosphere composition correspond-
ing to the current nominal atmosphere.
The heat shield thickness shown in Figure 6-3 is based on
2000°F backface temperature. A number of insulative approaches
can be used to reduce the temperatures below the 2000°F level.
For Saturn/Uranus, our baseline approach is to hollow-out the
carbon phenolic below the 2000°F isotherm whereas for the Jupiter
heat shield we have elected to forfeit the weight savings pro-
vided by the hollowed-out layer in order to increase the inherent
safety margin.
Figure 6-4 illustrates the similarity in entry heating and
pressure between planetary probe and mission flight entries. The
missile body point of interest is the control surface that was
protected with a carbon phenolic heat shield. Heating rates on
the missile nose tip are even higher but stagnation pressures are
sufficiently high (above 100 atmospheres) to exclude the appli-
cability of these data for planetary heat shield designs.
VI-5
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.i _¸'- .
•._L_:;_/ ._
,_ • ,e
'-:'L 7¢i!
• ,,,.7_-_ ', _
,,:'>_',_>,_;i_!C:__
7¸ _,, _;
1.8
1.6
1,4,
1.2-
1._ _
JUPITER
I I
-6 -8
'1 - DEG
NOMINAL ATM
-10
Figure
_L
X
T
06
0.5
::':.:'.::.;;;/.'ii1i! '_!';_'::;'.I': CARBONPHENOLIC
URANUS
I !0.7 i
I I
I
I
J
!
I
I
f --'----4
0.3 J fI I
L _ i I J
-20 -30 -40 -50
"YI -DEG
O.l
0.6
0.5-
0.4
0,3
-10
6-3. Heat Shield Requirements
SATURN
I
I
I
?. I
I
I
1
!
I I
I !
-20 -30 -40
YI - DEG
VI-6
: ,,, ".c
":L
.... j
J
-: L .
." !
, _i.!_/_,_ i___,_
=%
I
--.I
LL
I--"
L/,.J
-r-
I--"
z
24
16
8-
0,
0
Z0
I--
I
la.J
=: 10
w
e..
I=.
0
0
PLANETARY ENTRY
[NET HEAT FLUX TO SURFACE J
URANUS /
YI -
I
4 8 12
TIME - sec
I STAGNATIONPRESSUREI
24
1@
SATURN 8I
_'I = -40°
I 0
16 0
2OU,ANU SATU,N
PITER
4 8 12 16
TIME - sec
10
MISSILE REENTRY
NET HEAT FLUX TO SURFACE
FLAP-_
EARTH REENTRY
CONFIGURATION
4 8
TIME- sec
I LOCAL PRESSURE ]
I
12
X = 6.0IN.
X
1 , ..... !
4 B
TIME - sec
= 2.0IN.
12
Figure 6-4. Planetary Entry Environments
VI-7
_ __/_i___
i:_ .,_,. i_ ¸,
f
?
i !
I
r
i}
iii! i
i_'__!'_ii
The comparison in heating is in terms of net heating to the
surface; i.e., the reduction in heating due to ablation blowing
and hot wall correction has been applied. The comparison is
made in this manner since blowing greatly reduces the planetary
heat flux but only slightly affects the turbulent heating on the
=
flap. Furthermore, it is presumed that there is no effect in
material performance between convective and radiative heating
for Carbonaceous materials since the incident radiant energy is
absorbed on the surface. If one accepts this assumption, then
they could use the missile flap data to base the probe heat shield
design. Note, that the pressure levels between planetary and
missile entries compare favorably. Pressure is important since
mechanical erosion for carbon phenolic ablators has been corre-
lated in terms of this parameter.
Mechanical erosion representsthe greatest uncertainty in
predicting material performance during planetary entry. The
central question is how the material recedes, does it recede
primarily due to chemical reaction and sublimation (thermochem-
ical recession), processes that absorb large amounts of energy
per pound of material consumed; or is there a large fraction of
material removed by bits and pieces (mechanical erosion) result-
ing in a reduction of material effectiveness. Causes for mechan-
ical erosion have been attributed to preferential oxidation of
the binder, high surface temperatures with large temperature
gradients and high aerodynamic shear and large pressure gradients.
For lack of adequate analytical techniques, we have resorted to
empirical correlation of ground test or preferably flight data.
The correlation shown in Figure 6-5 is based on the missile flight
data discussed earlier. The correlation is in terms of measured
total recession rate, mechanical and thermochemical included,
ratioed to the predicted thermochemical recession rate versus
surface pressure and net heat flux to the surface.
A high degree of uncertainty is present in the application
of this correlation to planetary entries, primarily because of
VI-8
• -FLIGHT DATA FOR
CARBON PHENOLIC FLAP INAIR
(TOTAL RECESSION 0.6TO 0.82IN.)
SURFACENET
HEAT FLUX
Btu/ft2 - sec
I2,000
10,000
8,000
._j
¢. ;:'J
" .":L !'I
• ;:'* ,,
1 ,,. 1. _ 1. 1 I 1 I
0 2 4 6 8 10 12 14 16
SURFACE PRESSURE - ATM
Figure 6-5. Mechanical Erosion Correlation of Missile
Flight Data
the difference in environments. However, the correlation is
presumed to yield conservative estimates of mechanical erosion
since the aerodynamic shear levels were much higher than those
expected on the probe.
The Jupiter heat shield thickness based on computation of
the thermochemical and mechanical recession and insulation re-
quirements for an 800°F bondline temperature are illustrated in
Figure 6-6. Assuming a constant forebody ablative thickness and
adding the honeycomb and powder insulation weight results in an
aeroshell mass fraction of about 53 percent. Although this is a
relatively high weight penalty, it is within the probe weight
allotment.
VI-9
.... ;i
(
• " *_.5
;::'I
NOSECAP REQUIREMENTS
NOMINALATMOSPHERE
Yl = -7"50
TITITITrT
STRUCTURAL HONEY-
SANDWICH COMB
IlllIIlll
INSULATION
MECHANICAL EROSION
THERMO-
CHEMICAL
RECESSION
::_BiO_ili_!
:!_O L!_i!i!::
i:_:_:_:!:,,'.i::::::
I'
0.85IN.
0.45 IN.°
A -
0.35 IN.
T
1.30 IN. '
'COMPUTED BY
Figure 6-6.
i_. 35.0i DIA _I
AEROTHERM CORP.
Carbon Phenol ic
FOREBODYHEATSHIELDWEIGHT = 176LB
FRACTION OF PROBEENTRY WEIGHT = .53
Heatshield for Jupiter Entry
A plasma jet test program •(Figure 6-7) was conducted in the
MDRL i0 Megawatt Facility to obtain performance data on carbon
.... 10MWMDRLFACILITY
NOSETIP MODELS WEDGEMODELS ENTRY FLIGHT PROBE
q CONV= 8000 Btu/ft2-sec
P = 10 AND 20ATM
H = 5000Btu/Ib
AIR AND N2
qCONV = 3600Btu/ft2-sec qRAD = 5 TO 45,000
P = 10ATM P = 2TOi5ATM
H = 3400Btu/Ib H = 2 x 105TO 5 x 105 Btu/Ib
AIR AND N2 H2/He ATMOSPHERE
0_ ° CLOTH ORIENTATION
0.30 INCH
PLASMA ,_,__..,, _._
_'/_'"--20_LAY UP
SA_4EAS
FLIGHT HEATSHIELD
5°& 450WEDGE
"_-90° 0 RIENTATION
RADIATIVE.
HEATING
Figure 6-7. Plasma Jet Test Program
VI-10
!phenolic at the possible highest heating rates but at moderate
pressures between I0 and 20 atmospheres. A key objective was to
evaluate the mechanical erosion phenomena in an oxidizing (air)
and an inert environment for possible extrapolation to the hydro-
gen helium planetary atmospheres. Both nose tip and wedge models
were tested in air and nitrogen plasma streams. Much higher
heating rates are feasible with the nose tip model, however, the
wedge model besides providing a larger test specimen, is also
more representative of the flight heat shield in regards to the
cloth orientation with the boundary layer flow.
Theoretical ablation predictions have been made to correlate
the measured recession rate data. As shown in Figure 6-8, a fair
degree of matching the data was achieved in our initial analy-
tical effort and work is continuing in this area to resolve some
of the discrepancies. A major problem is the uncertainty in the
nose tip recession rate measurements. Contributors to the uncer-
tainties are the relatively small total recession experienced,
the initial swelling of the material and the lack of sufficient
data points to provide a good average value. Recession measure-
ments were obtained from measurements of the before and after
test specimen thickness and from motion picture views of the re-
ceding surface. The nose tip motion pictures showed small flakes
of carbon phenolic laminates being removed (mechanical erosion)
in both the air and nitrogen runs but at a higher rate in air
tests. The small nose tip size and the flat laminate lay-up con-
tributed to this mechanism of removal.
Although a number of discrepancies are indicated by the data,
the trend of the data indicates a higher mechanical erosion in
air than in nitrogen and higher erosion rates in the turbulent
higher shear wedge environments.
MR. VOJVODICH: Sam, I think this will probably be a ques-
tion of general interest, and that is: In the Saturn and Uranus
cases you show, as you decrease the entry angle, the heat shield
VI-11
GAS CLOTH COLD WALL
STREAM ORIENTATION . HEAT FLUX
(DEG) Btu/ft2-sec
SURFACE MEASURED PREDICTED
PRESSURE RECESSION THERMOCHEMICAL MEASURED
ATM RATE RATE THEORY
(INCH/SEC) (INCH/SEC)
• - ,r . i +
m _ ..... _
: =+_;_?;"i
+,_+.., _; , _;+, +
++ +
- . . ++.
NOSETIP DATA (LAMINAR)
AIR 0 6300" 10 0.058 0.057 1.01
N2 0 6300" 10 0.025 0,035 0./2
AIR 0 6800" 20 0.087 0.070 t.14
N2 0 6800' 20 0,022 0,031 0.70
,u+,,'_ +. +'
:'_: -+2".<.
•... _-.,+.+
,,+. +}
• . +.+ i
+ • "+ i
:, +
AIR 20
N2 20
WEDGE DATA (TURBULENT)
4560°° 13 0.064
2630 13 (, 0.005)
(SWELLING)
0.050 1.28
0.006
•. (
'HEAT FLUX MODIFIED BY NOSE TIP SHAPE CHANGE
"'HEAT FLUX BASED ON CUSPED CALORIMETER DATA
Figure 6- 8. Preliminary Comparison of Plasma Jet Test
Data with Theory
i!
.+...,.m":.? :+'_
". ;..
• k, :.7
_ :.+-_+? 0, .:
+ _'P'_:.+'?; i _:'i
•'+2+'c_.'i_?::,,,i
++5::++++;+'.'++,+
: +C • •::".? ',
weight goes up. In the case of Jupiter, as you are decreasing
the entry angle, the heat shield weight is going down; will you
comment on that.
MR. MEZINES: The total heat shield thickness is the sum of
the recession thickness plus the insulation thickness needed to
limit the backface temperature to a certain value. In general,
increasing entry angles result in higher recession but lower in-
sulation requirements. The total thickness or the sum of these two
thickness may or may not increase with higher entry angles but will
depend on which mechanism predominates. For Jupiter entries, re-
cession is the dominant mode, thus total thickness requirements
are higher with increasing entry angles. Conversely, for the Sa-
turn/Uranus entries, the insulation requirement sizes the
VI-12
heat shield thickness; thus higher entry angle entries require
less thickness to achieve the same backface temperature.
DR. NACHTSHEIM: Our next speaker is John Lundell who will
describe the evaluation of graphitic materials in the Ames high-
powered gas dynamic laser.
[
VI-13


Carbon phenolic heat shields for Jupiter/Saturn/Uranus entry probes

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