The characteristics of the Mars entry including the mission sequence of events and associated spacecraft weights are described along with the Viking spacecraft. Test data are presented for the aerodynamic characteristics of the entry vehicle showing trimmed alpha, drag coefficient, and trimmed lift to drag ratio versus Mach number; the damping characteristics of the entry configuration; the angle of attack time history of Viking entries; stagnation heating and pressure time histories; and the aeroshell heating distribution as obtained in tests run in a shock tunnel for various gases. Flight tests which demonstrate the aerodynamic separation of the full-scale aeroshell and the flying qualities of the entry configuration in an uncontrolled mode are documented. Design values selected for the heat protection system based on the test data and analysis performed are presented

Polutchko, R. J.

English

NASA Technical Reports Server

_ !_?i_i_
r,
L
.... i
VIKING ENTRY AERODYNAMICS AND HEATING
Robert J. Polutchko
Martin-Marietta Corporation
MR. POLUTCHKO: Entry into the relatively thin Mars atmosphere
is pretty straightforward compared to some of the more exotic out-
er planet entries you have been hearing about. Figure 5-20 des-
cribes the characteristics of the Mars entry including the mission
sequence of events and associated spacecraft weights.
The Viking spacecraft is comprised of a modified Mariner
Orbiter and the Viking Lander Capsule. The Mars Orbit insertion
weight is 5189 pounds. After separation of the entry vehicle,
the de-orbit maneuver is performed by a low thrust, long burn
time (15 minutes) propulsive maneuver. This propulsion system
is a mono-propellant hydrazine system that is also used for re-
action control during entry. During the coast period (3 to 6
hours) after de-orbit, the entry vehicle is oriented to an angle
of attack of -20 degrees in order to align several entry experi-
ments with the free-stream velocity vector. I will describe the
locations of the entry science sensors in a moment.
Atmospheric entry is arbitrarily defined as 800,000 feet and
the entry vehicle weight is 2060 pounds. At 0.05 G's decelera-
tion the entry vehicle reaction control is switched from pitch,
yaw and roll attitude hold into a rate damping mode for pitch
and yaw. The Viking entry vehicle flies a lifting trajectory so
roll attitude hold is maintained to control the lift vector.
Parachute deployment is provided by the guidance and control
system radar altimeter at 24,900 feet. Depending upon the at-
mosphere encountered the mortar fire Mach number will be between
0.6 and 2.1. The aeroshell/heat shield is aerodynamically sepa-
rated 7.0 seconds after mortar fire. The terminal propulsion
engines are ignited at 3565 feet above the surface and the para-
chute and base cover are separated 2.0 seconds after engine start.
The terminal propulsion system is also mono-propellant hydrazine
and the engines are differentially throttled for pitch and yaw
V-47
https://ntrs.nasa.gov/search.jsp?R=19750012311 2020-03-19T21:09:41+00:00Z
• > _:i•'<_
.i
i
_ !
I
Z
(2)'
_D
Z
Z
-r-
D
o
e_
"r"
%--
>-
I---
Z
Z
n
m
0_ _ 0
A
v
0
0
o ¢-"
._
e- i
¢M
ORIGINAL PAGE IS
OF p_-_ quALr1_-
V-48
o
('4
!
u%
D_
control. Roll control is provided by small roll engines mounted
on the terminal propellant tanks. A constant velocity descent
contour is reached above the Mars surface and the Lander engines
are cut-off at surface contact. The touchdown velocity will be
approximately 8.0 feet/second.
The Viking entry into a relatively thin atmosphere is criti-
cally dependent upon high drag. The configuration as shown in
Figure 5-21 is a 140-degree included angle cone with a base cover.
There was, of course, considerable concern with the aerodynamic
stability of very high drag configurations but we will discuss
the stability characteristics in more detail later. The entry
configuration is eleven and one-half feet in diameter. On the
windward meridian several entry science instruments are located -
an upper atmospheric mass spectrometer, a retarding potential
analyzer and the stagnation pressure port. A stagnation (recov-
ery) temperature sensor is located on the leeward meridian and is
deployed through the heat shield at a velocity of i.i km/second
(Mach 4.0). We also have some engineering measurements located
on the heat shield (four diametrically opposed pressure ports) and
one base cover pressure port.
7
.;; .l
:" : 7 i
Sometimes the more simple points are overlooked. For a very
blunt vehicle lift is obtained from the high axial force. The
body force diagram is shown in Figure 5-22. In order to obtain
a positive lift from the axial force, a negative angle of attack
is required. The normal force is also negative but is a small con-
tributor to the resultant lift vector. For the Viking configura-
tion the lift to drag ratio is given approximately by -0.015e. For
a c.g. offset of -1.84 inches the trim angle of attack is -11.2
degrees and the L/D is 0.18.
Figure 5-23 presents test data for the aerodynamic character-
istics of the entry vehicle showing trimmed alpha, drag coeffi-
cient and trimmed lift to drag ratio versus Mach number. The MD
requirements here refer to the mission definition requirements for
V-49
"--'-" :'_:i_¸I
_ _'r --, "_"
r
 i((iii:i,,
. °
• r
._ " ,
t
t "
t__ ,:;.:_!_
Z
0
I--
m
¢.)
Z
0
m
1.1_
Z
0
¢.)
¢.)
m
-i"
I---
Z
1.1.1
/
O0
\
\
\
V-50 '
+
N
c-I
c-
o
o--
E
I--
,-'4
C'q
I
L_
_J
-,-I
_!_ i__':'I_
!il_-_i_:i_:t
O,S_I:,:_!
:_.,._!i:_:
.\
,I
*; : _=
un
I.L.I
¢..,)
o
>-
0
:>
>-
I..,.,-
Z
C_'< (..)'-_
\\ \
V-51
u'%
c_
I
0
o
_-" w
v_ (.2
II
I!
"'3 '_
O
0m f"
| 4-
ozoz
¢N
¢N
!
Im
-,-t
._._ i _q
t_
: i
:::tq
:. T _]
,:., ,
f
•_%1_!_!_ _,_
,.., 0 ,,_ @,J (_ _ ,,
"_ _ _ _ o _ ,\
_o <5 _5 o d d ",
I r-- I r - _: u'_ ' t.... i I I "_-_"_"_.-.
I I I I I i _ [--
I I I i I I I , _ _
I i r 1 I I _ ._o i
/ I , J _i , I < _2 !o
J , I i I i I _ ._ _ -I_
/ '_! , -._i i '-" I ' E _-_: I
/ I I I ! I I "__. _ " I ,,--_
I \ , _, I ' _ ,,, ,_ .c::: I -
I ,,i--_l-,',, "\_ _, I e I . _ _ I
I 'C_..'_ _"_ _ _.. _ H I I°
/ / / ,/ ./ / / I-
I / / / I I :_
/I _1 __1 I I Io=
t) l " lel , e, -
' --'"--- C_+."" '----_ _-. ,_-. I"-'
/ %_Z_, "_.,_, "-__ _. "1_
I "-' m \ _. " \. _, \ \ I
_' /2,,,', ,_ ii _ , I
/- L.,,,,,'ll IA: I .A. I l'_::rI¢-_ -_-_ ,_ !_ _:_I_ -J,.._/> ! , , , ', I L-
k ! I I _ i.l. I I l _i i __t"
• _ _ _ 1"_ _._
V-52
¢'o
¢xl
I
-M
E:,_,I
I
:. }
> ',i
,-_ ._ ..._.'; ::_
:Lt:?,f _. / ::::.t:_
!:_':,t!%<-_,
atmospheric reconstruction. The specification requires a priori
aerodynamic coefficients within + 5% and the test data certainly
falls within the indicated tolerance. These test data were ob-
tained using conventional wind tunnels and fairly straightforward
testing technology.
Figure 5-24 shows the damping characteristics of the entry
configuration. These data were experimentally derived utilizing
forced oscillation and free oscillation testing techniques. This
figure shows the basic negative damping at low angles of attack
for very blunt configurations. The plots of Cmq plus Cme versus
and the same parameter versus Mach number show that there
are two Mach numbers (about 1.2 and 2.0) where we have negative
damping at low angles of attack. It should be noted, however,
that for a trim angle of attack of -I!.0 degrees that the Viking
configuration has positive aerodynamic damping at all Mach num-
bers. Also note the relative insensitivity of longitudinal c.g.
position on the pitch damping values.
A1 Seiff (NASA/ARC) is currently in the process of obtaining
ballistics range (free flight) test data for the Viking configu-
ration. Comparisons of foreced and free oscillation data with
the free flight data Should provide additional assurance of the
predicted vehicle motions.
On Figure 5-25 the angle of attack time history is shown for
several Viking entries. Again the entry altitude is defined as
800,000 feet above the mean surface level. As I mentioned earlier,
the nominal trim angle of attack is -11.2 degrees when Viking en-
ters the sensible atmosphere. At the end of the long coast period
following the de-orbit maneuver the guidance and control uncer-
tainty (worst case) in angle of attack is _ i0 degrees. For en-
try science reasons we have a pre-programmed attitude hold mode
prior to entry into the atmosphere. The angle of attack will be
-20 degrees which orients the windward meridian directly normal
to the velocity vector for the mass spectrometer and RPA data.
In the worst case then, alpha could be either -30 degrees or close
V-53
|.,%
\
t,
"_J_ ..,.-, C"_¸ t
_,'_ '_
..¢:._ _=" _..
=,,o_
_. ---.._.
'.,/-54
ORIGINAL PAGE IS
OF POOR QUALIT_
'_:_L_._i_'_
Z
• i
!
1
1
I
I
l
1
I
+1
\
\
II
I !
! I
V-55
U%
!
U'_
(D
tm
-,-I
[.,.,
to the trim angle.
-30 degree case.
Our discussion here will be limited to the
r
i_
A normal gravity turn will change the angle of attack as in-
dicated. At 0.05 G's we switch to rate damping and combined
with the natural aerodynamic damping characteristics the vehicle
motion rapidly converges to the trim alpha. Shown _ on this fig-
ure are two atmospheric extremes and the convergence associated
with only natural aerodynamic damping (i.e., reaction control sys-
tem inoperative). It should also be noted that the reaction con-
trol system is operating in opposition to the aerodynamic damping
forces in order to maintain the pre-pr_grammed angle of attack.
These engines are 4 pounds of thrust each (4 engines). After
reaching the trim angle of attack maximum excursions due to spec
gust profiles (20 meters per second) show maximum excursions of
3 degrees to vehicle attitude.
Figure 5-26 presents the relatively mild stagnation heating
and pressure time histories. The curves are the worst case de-
sign limit values and represent atmospheric, entry angle and
lift to drag ratio extremes. The stagnation heating values are
calculated using a Newtonian pressure gradient and the Marvin
and Pope correlation with real gas effects included. This rela-
tively mild environment allows us to use very lightweight struc-
tures and heat protection and, therefore, the normal care of
design and test must be exercised to provide a minimum weight
entry vehicle.
Figure 5-27 presents the aeroshell heating distribution as
obtained in tests run in the NASA Ames 42-inch Shock Tunnel for
various gases. We also have obtained equivalent data in CF 4 at
NASA Langley and in air at Cornell. The solid curves are our
predictions of a heating distribution using the Aerotherm BLIMP
C program. All our data and predictions have correlated quite well
and an example of the agreement is given here. This high heating
rates at the corner of the aeroshell are caused in part by the
sharp radius - 1 inch full-scale. The differences indicated
V-56
n: z
=,
.:i
: !
:s
:: :
i_i__:::
: i_.-_:I_:
Q_
0
m
-I-
£.9
Z
m
-1-
Z
<D
m
z
<=.9
.<
(/I
C3
Z
m-"
Q.
C.)
2
>-
I I I I
,-4
\
\
c;"
V-57
I
%
0
0
c_
cq
I
D_
.,-4
.. ; ,7 _-
• .,J ; -.
. -_!:.: i;i_,_,;_
cS c_ _ ¢S
V-58
o,1
I
-,-I
\between the BLIMP C prediction and the data is test model pecu-
liar. We have obtained data on a model constructed to emphasize
specifically the instrumenting of the sharp corner. These data
indicate that the BLIMP predictions shown here are accurate.
These predictions here are based on the pressure distribution
data from that special model and the heating rates indicated by
the test data shown here are, in fact, in error.
On Figure 5-28 is presented some heating data from the Variable
Density Tunnel at Langley at Mach 8.0 in air. Also shown are
BLIMP laminer and turbulent heating rate predictions. The lee-
ward side of the aeroshell seems to experience a transition to
turbulence at Reynolds numbers between 3 and 4 million. We ar-
tificially tripped the boundary layer and experienced additional
increases in the local heating rates which seem to show a good
resemblance to the turbulent predictions. The Viking Reynolds
number at the peak heating point in the worst case trajectory is
about 3 million and the evidence seems to indicate that we could
expect transition on the leeward side. This Reynolds number
translates to a momentum thickness Reynolds number of about 140.
_ _! ..i _
Precise transit criteria is not the point here since many fac-
tors influence determination of such a specification. However,
this wind tunnel test, in fact, was a very close flight simulation
for Viking and in the same facility Apollo tests showed remark-
able correlation with flight test data. The Viking heat shield
was designed to handle the situation indicated by these data.
We also placed the entry science recovery temperature sensor on
the leeward meridian to take advantage of the higher local Reyn-
olds numbers at that location.
The curve of Figure 5-29 presents the design values selected
for the heat protection system based upon all the test data and
analyses we have performed. Basically, we have taken a conserva-
tive approach that calculates the expected heating rates in the
V-59
r •......
:T_L_!_/,¸,_
7-,, .:ii
fi . .'
. , _7,1
-7.[- _:
, . . -_
. . . .,._
_b,7-.__._
I!
0
,,> = __._; _/¢,_ + u + + n
- obooo
_ --.J
e_
S
I
"o o\
_o
V-60
0
_3
d
c5_
c_
0
c_
I
c_
I
I
130
I
.jk
,1
L.-_
." ...
_,, i, .¸_,_
?_i_, ii!ii_!_i
Z
o
i
I
i
Z
u
Z
I
Z
I
I
o
o o 0
II II II
V-61
x
0
Z
i
0
_0 __1
¢M
0
0
¢-
0
fXl
I
U'3
_ LS ,_i_,p_
_!i_!_,_i_I_'_
_ii_ _,_,_!_i_
¸if!•¸
• ir_rll¸ Li
p,
_i•
__i_ __ _
Mars CO 2 atmosphere by using a measured freon pressure distribu-
tion. The shock density ratio basically governs the pressures
and the values for freon and CO 2 are very similar. For the tur-
bulent areas we have modified the heating values using BLIMP
rather than, for example, the Harris model at LRC. BLIMP gives
a factor of three increase in this area while the Harris model
shows about a factor of two. The stagnation area does not really
experience a "Newtonian" stagnation heating rate but we have used
the full stagnation value for design.
Figure 5-30 shows some test data we obtained on protuberances.
The case shown here is the mass spectrometer cap which is poten-
tially the largest if it failed to jettison prior to entry. The
interference factor above the local "smooth" heating rate is
plotted versus streamline direction. It can be seen that a fac-
tor of about 3.0 increase in heating rate could be expected. We
have locally protected these areas with a high density ablative
material that was previously flown on the USAF PRIME vehicle.
Figure 5-31 presents the real gas effects on the entry vehicle
aerodynamics based on CF 4 data we measured at NASA-LRC and some
preliminary data measured at NASA-ARC. You will note the slight
increase in drag and the more non-linear nature of the pitching
moment with alpha. However, the trim angle of attack for all
three test gases is virtually the same for the Viking configuration
at -11.2 degrees and the lift to drag ratio is virtually identi-
cal. We don't anticipate any problems for the lifting entry aero-
dynamic performance in the Mars atmosphere.
Figure 5-32 summarizes several of the design values and design
factors for the Viking entry mission. The heat shield is basic-
ally an insulator and is, therefore, total heat rather than heat-
ing rate sensitive. The base cover is designed for 2 percent of
stagnation heating based upon test data. The maximum base cover
heating rate that was measured was 1.5 percent of stagnation. We
have applied a design factor of 1.5 to all heating rates for
smooth areas and a factor of 4.0 to all protuberances areas. Shear
V-62
_ ?i'_¸ ,_._._,:
:.:;__'_,_i i
_ "_i_: _.i_7
7
. . ':'7
!
N
+
n
o t-
¢_1 t-
p__ I--
o
/ _D
o _g_
0
u
._ |
I I I !
.._1
V-63
oo
.=_
h_
C'4
0
0
I
un
,:E>,_::i:_:!il
_:_:;_,i::_i•5,
. !
• :J
L_
C.)
m
Z
>-
O
m,-
Z
O
u,%
O
w
(.D
.<
.<
Z
m
£2-
X
I,
I, I ! 1 I __I
I",-..
_._I
(.D
0
,...i
V-64
I I I C_
K..
Q_
o
_-" -r-
C3
O
Z
_D
_D
-oO_
O
_D
m
O
.i)
,-4
e'%
I
U3
_m
: . :',
J
• • k. ;
_ _i!._ ?i_;_!
0
I,,,-,,,
Z
¢,.,0
M
Z
<>
I
I.I.J
0
L.I.J
,-'4000 I"_- U'%
o
e/,_ o
(2) !,._,
..e.o_ _.
I
H U II
._s'
ZI=
0
LL!
A
_ =_ _ _L_ _
,,,: _ _ _. __.
__ _._ _
<_ ..__
V-65
(-q
!
u_
-,-4
. _ "- , 2
' . ;L i
- " ;. _
_.':-_.. • i
stress factor is 1.5 and aerodynamic loads factor is 1.25. These
factors are applied to worst case combination of atmosphere model,
entry angles and lift to drag ratios.
I would now like to show you a five minute film clip of the
qualification flight test program of the Viking decelerator sys-
tem, the Balloon Launched Decelerator Tests, BLDT. As summarized
on Figure 5-33, the program consisted of four tests conducted at
the White Sands Missile Range (WSMR) in New Mexico and were de-
signed to span the extremes of the worst case conditions on Mars.
These flight tests also demonstrated the aerodynamic separation
of the full-scale aeroshell and the flying qualities of the entry
configuration in an uncontrolled mode.
The parachute is a disk-gap-band configuration 55 feet in dia-
meter, mortar deployed in a single stage with a mortar ejection
velocity of about i00 feet per second. Tests were conducted at
Mach numbers of 2.2, 1.2 and 0.5 and dynamic pressures of 14.5
and 4.5 pounds per square foot. The full-scale Viking test ve-
hicle was carried to 120,000 feet by a helium filled, 34 million
cubic foot balloon. The test vehicle was dropped from the bal-
loon and rocket boosted to the test altitude and Mach number. All
tests were successful and demonstrated a 35% structural margin
above the worst case expected at Mars.
V-66
L<:',._.__',,_:
_ :2 L r T
,r- q
: .. ?
.I,
i, '" _i:¸ §
e_
e--
cO
,m
¢0
i...
_:>
0
._ _. o_
_ =
,-, #. _ <
0
,..[
V-67
M
I
u_
D_
-,4


Viking entry aerodynamics and heating

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server