Fellenz, D. W.

English

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RECOVERY OF ORBITAL sTAGES
By
Dietrich W. F ellenz*
6 5 8. 3
The reasons to be interested in the recovery of a stage that
reaches orbital injection conditions (usually a secop._ stage) are
basieallythe same as for the recovery any other l_iece of space hard-
ware:
1. Post-flight inspection affords the detection of desi&n short-
comings and a better evaluation of the actual environment of the com'-
ponent (loads, heat input etc.).
2. Reduction of cost per pound of payload in orbit due _o re-use
of hardware.
3. Operational advantages of positive c_spOS_ Of hardware
and if possible return to the refurbishment _ud l&_nch site.
Wh_le post flight inspection is always desir_le from an engineers
point of view in order to advance the state _ the art, it lookm like that
the development of a recovery system can only be so_l on the basis _f
points Z or 3 above.
To prove the desirability of recovery on a cost basis alone
would require that all developmental and operational costs referred t_
the reduced payload in orbit would come out cheaper than in the case
of an expendable reference vehicle. St_ies performed o_- contracted
by h4SFC in this area showed that this point could be proven for first
stages assuming the present state of the art. The _iscussion of cross-
over points, of course, is influenced very strongly by the basic coat
assumptions. At the present time, it seems, that no cost reductions
can be derived from second stage recovery.
The third and by no means less important aspect is the oper&tional.
It can he expected that the volume of launch operations in support of
Division, Marshall Space Flight Center
https://ntrs.nasa.gov/search.jsp?R=19730061698 2020-03-17T13:47:51+00:00Z
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orbital operations, lunar and planetary missions will continue to grow
and will reach dimensions where the controlled disposal of all spent
space hardware will become mandatory. Taking an expendable vehicle
with such a "disposal" system and its reduced performance as reference,
it r_ight prove that fuii recovery and return of all stages can become
economical. The requirements for recovery forces will grow pro-
portionally to the volume of the launch operations. It is Obvious that
the capability to return to the launch base has to be more and more in-
corporated in the vehicle. This would in turn speed up the refurbish-
ment and increase the overall flexibility of the operation.
That means that the first stage requires sufficient propulsion
for fly-back, and that the second stage glides back to the launch site
a_ter one or more revolutions around the earth and subsequent aero-
dynamic re ,entry.
To Study the sensitivity of various parameters of recovery the
Marshal Space Flight Center sponsored three industry study contracts
(HAS 8-1513/1514/1515) on the subject "Study of a Two to Three Million
Pound Thrust Launch Vehicle". The basic mission was defined as two-
stage to 307 N. h_. orbit. Recovery was to be considered for both stages.
Fig. I shows a typical mission profile.
An evaluation of the final reports of the three studies with respect
to structural weight increases due to recovery was made and the re-
sults are shown in Fig. Z. The parameter shown is the weight of the
recovery system in percent of the structural weight of the expendable
reference vehicle, based on equal propellant ratio, i. e., on equal
ideal velocity increment of recoverable and expendable stage. The data
generated by the different companies scatter considerably. This is
partly due to the different assumptions with respect to structural
efficiency as indicated by the structure ratio of the expendable reference
vehicle shown in Fig, 3, partly due to the relative novelty of a
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particular recovery mode. We expect 'tO be able tc_-sa'nooth out some of
the scatter in these data after a presently going study of fixed wing
recovery systems .has been evaluated, In order to get a better feel
for the performance penalty associated with orbital stage recovery by
paraglider a conceptual design study was performed at MSFC, the
results of which will be discussed later in some detail.
In the case of a two stage to orbit configuration we find that
there is a payload decrease of about 1 ib per 5 Ibs increase in first
stage structure weight and a payload decrease of i Ib per i Ib _crease
in second stage structure weight.
In addition to that the increase in second stage structure weight
due to recovery is considerably higher than that for first stage recovery.
This is mostly so because of the more severe re-entry environment
and the much longer glide and exposure times requiring heavier thermal
protection.
This explains why second stage recovery is so expensive in terms
of payload. Fig. 4 shows the effect of second stage recovery on the
payload of a two stage to 307 N. IvY. orbit configuration with an initial
weight of 2.4.106 lb and 3.106 lb thrust. First stage LOX/RP; Second
Stage LOX/I._I z. The ascent trajectories utilized intermediate l_rking
orbits and H0hmann transfer up to 307 N. IV[. altitude. The recovery
factor, as defined by NA.Alsee Fig. 4 for equation,_ represents the
ratio between the stage structure weight factors of the recoverable a_d
the expendable reference vehicles. The figure shows on its left side
for K z = 1.0, which means no weight added for second stage recovery,
the payload performance of the corresponding lower stage (again with
or without recovery) carrying an expendable second stage.
Some of the scatter in the payloads shown can be explained by
different staging orbit altitudes and different "kicker _systems" to
*Conceptual Design Study of Ten Ton Reusable Orbital Carrier
Vehicle• NAS 8-_687/5037.
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perform the transfer maneuver up to the target orbit. The recovery
modes suggested ¢or _+udy,_ +_'_ "Z-3 _a.'_.-...... " .... _,,,_,_on Pound Thrust Launch
Vehicle Study" were "Paraglider" or "Fixed Wing". The lightest of
these modes of course is the Paraglider, although, as you saw from
Fig. 2, this sytem can amount to a sizable weight penalty. Increasing
second stage recovery factor Kz means heavier and more sophisticated
recovery systems, usually associated with extended cruise capability.
I would now like to present some details on our parametric
design study of the application of a paraglider to the recovery of an
orbital stage.
The paraglider concept looked attractive to us because of its
light weight, the simplicity of the system, the possibility to stowe
it away in a fairly small volume along the stage which would not penalize
the vehicle configuration during ascent, and the inherent stability of the
paraglide r configuration.
With respect to the mission we assumed that the payload shall
be delivered in a 307 N. M. orbit using a two-stage plus "kicker-
stage" arrangement. The second stage burns out at low altitude at
a velocity equal to the local orbital velocity plus the velocity increment
for Hohmann transfer up to 307 N. M. Then it was assumed that the
empty stage plus payload were injected into orbit. After waiting in
orbit the orbital stage was brought to re-enter with a zero altitude
virtual perigee, corresponding in this case to a flight path angle of 9_
deg at 400,000 ft altitude.
Starting from this condition we investigated the influence of
paraglider wing loading and deployment altitude on the thermal protection
requirements and the overall structural weight of the paraglider package.
The characteristics of the stage were those of an early version of a
Saturn second stage.
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The wing loadings considered were 1. Z5; 5; 10 lbs/ft z. The
deployment conditions investigated were 400,000 ft altitude; maximum
dynamic pressure, and finally Mach 5.
Upon entry into the sensible atmosphere a drag device would
be deployed to stabilize the stage. This drag device would be retained
after deployment of the parawing. The de-reefing of the wing was
controlled to keep the normal acceleration of the stage below a certain
limit. The following assumptions were made on the part of the paraglider
system:
I The physical dimensions of the paraglider wing installations of
different wing loadings are assumed to be geometrically similar;
• Keel length equals leading edge length for easy stowing;
• Wing leading edge sweep angle in fully deployed condition is _= 50 °
* C. G. location required to fly at subsonic L/Dma x and 11% static
margin is 0.65 _'below wing leading edge, and 0.55 _'behind leading
edge of _;
• The wing would be oriented at an angle of attack that yielded max.
L/D for that particular wing/body combination; supersonic flow:
u _40°; Subsonic flow: _Z5°;
• The stage body is always oriented parallel to the flight path;
• The net structure weight of the stage, which is equal to the weight
recovered was W n = 41,000 lb;
• The basic structure weights of the paraglider packages were obtained
by scaling with respect to wing loading;
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w/s
[lb/ft
15
I0
5
I.Z5
Ws s /W n
[%]
13
16
Z5
78
Assumes load factor
n:6
In scaling of the structural weights from a 15 lb/ft z wing loading base
point vehicle the following assumptions were made:
/1. Wing structure weights scale i proportion_al to wing area,
i. e., inversely proportional to wing loading.
:' ly
Z. Cable weights scale inversely proportional to the square
root of the wing loading under the assumption of a geometrically
I_,_;lar suspension systen-,. _,=y length affected. Loads and _ are
same. )
3. Landing gear, control system and drogue body structural
weights are roughly independent of wing loading.
We ran re-entry trajectories deploying wings of the different
wing loadings at the different points along the trajectory. The results
of these runs were fed into a thermodynamic analysis to determine the
heat protection required. It was arbitrarily decided to use an ablative
system. The basic stage structural material was changed from Aluminum
Z014 to stainless steel.
The ablation material weights were then determined, added to
the glider structural weight and referred to the net structural weight
of the recovered stage. The results are shown in Fig. 5.
In this figure it is considered that in the cases of deployment
at 400,000 ft altitude the maximum resultant load factor almost in-
dependently of wing loading was not higher than 3 g_s, and that in the
cases of deployment at qrnax and iVIach 5, the max. resultznt load factor
incurred was I0 and 9 g_s respectively. The weight of the glider was
then adjusted assuming that the structural weight scales directly pro-
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portional to the load factor.
The main trend of the curves on Fig. 5 seems to indicate an
advantage in going to higher wing loadings, i.e. , smaller wings. Further-
more the curves would indicate a preference for deployment ,at 400,000
ft altitude. However, there is a design difficulty in that it is hardly
conceivable how the suspension cables with a diameter of in the order of
_ Normal load factor d_ploymen_s
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Z in, and an additional ablation coatlng of in the ;order of 1/7 in. could
be stowed and then deployed within a split second without loosing the
ablation coating. No such coating is required for the lower altitude
deployments.
Therefore, our tentative conclusion at this time is to prefer
to deploy the wing below Mach 5, preferably at subsonic speeds and
to go to as high wing loadings as are compatible with the overall flight
stability and glide capability to ensure safe automatic landings. We
feel that even the application of a radiative cooling system for the case
of deployment at 400,000 ft altitude would not change this preference.
If the subsonic glide capability of a paraglider is not required, a very
similar system can be based on a parachute. The resulting weight
penalty would be very low but has to be bought at the expense of
impact and retrieval problems.
At the present time it cannot be stated positively that orbital
stage recovery will save costs, however it can be said that from the
operational point of view it would be very attractive. Advances in the
state of the art of recovery systems will reduce the weight penalty
associated with reusability, and in general will tend to make orbital
stage recovery also attractive from the economical aspect.
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COAST / TRAJECT.
TRAJEC' _ RBITAL STAGE
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BURNOUT OF ORBITAL
STAGE
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TYPICAL ORBITAL MISSION PROFILE INCLUDING RECOVERY OF T}_ ORBITAL
STAGE IMMEDIATELY AFTER ONE REVOLUTION AROUND T}LE EARTII.
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Recovery of orbital stages

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