nasa cr 112205
TRANSCRIPT
NASA CR 112205
N 7 3 - 1 3 9 0 9
DESIGN
OF A
SPACE SHUTTLE
STRUCTURAL DYNAMICS MDDEL
Prepared Under Contract NAS 1-10635-H by
Grumman Aerospace Corporation
Bethpage, N.Y.
For
NATIONAL AERONAUTICS AND SPACE ADMINISTRATION •
ABSTRACT
A l/8th scale Structural Dynamics model of a parallel burn
Space Shuttle has been designed. Basic objectives were to re-
present the significant low frequency structural dynamic
characteristics while keeping the fabrication costs low.
The model was derived from the proposed Grumman Design 619
Space Shuttle. The design includes an Orbiter, two Solid Rocket
Motors (SRM) and an External Tank (ET). The ET consists of a
monocoque !£>„ tank (.02" walls and .016" lower dome), an intertank 1
skirt (.05" skills) with three frames to accept SRM attachment
members, an LH tank (.025" and .016" skins) with 10 frames of
which 3 provide for orbiter attachment members, and an aft
skirt with one frame to provide for aft SRM attachment members.
The frames designed for the SRM attachments are fitted with trans-
verse struts to take symmetric loads. The SRM consists of
a monocoque (0.2" skins) cylinder representing the propellant
carrying structure, a simplified forward skirt with two frames
and longerons for interstage attachments, and an aft conical
skirt with one frame for interstage attachments and U longerons
representing the on-pad support structure. The orbiter consists
of an aft section representing simplified version of the major
load paths between engines and the aft interstage attachment,
a midsection formed from U shaped frames spaced about 10" apart
covered by a ,02:I skin, provisions for 3 different payload
lengths, wings designed as 6 spars covered by .02" skins, a
simple torque 'sox representing the fin out to the e.g., and a
tapered non-circular shell representing the cabin.
The model design details are presented in Ul drawings
which have been filed with the Dynamic Loads Branch of the Loads
Division at the KASA/Langley Research Center.
FOREWORD
The work described in this report was performed by Gruraman
Aerospace Corporation, Bethpage, New York, under NASA Contract
NAS1-10635-11 and administered by the Dynamic Loads Branch,
Loads Division, NASA/Langley Research Center, Hampton, Virginia.
The reported work was carried out between June and October 1972.
The design work was performed by A. P. LaValle, P. W. Tracy,
F. L. Halfen, C. M. Cacho-Negrete, P. J. Cartensen and
S. Rosenstein. The stress analysis was conducted by W. P. Bierds.
iThe suspension system frequency analysis was performed by
L. Mitchell.
This project was directed by M. Bernstein as one of the
Master Agreement Tasks Program managed by E. F. Baird.,.r
The guidance and assistance of S. A. Leadbetter and
U. J. Blanchard of the NASA/Langley Research Center is gratefully,i
acknowledged.
ii
TABLE OF CONTENTS
SUMMARY 1
MODEL DESCRIPTION 5
External Tank 5
Solid Rocket Motor •. 9
Orbiter 11v
SUSPENSION SYSTEM FOR MODEL 1^
Suspension Concepts 1^
Recommended Suspension Scheme lU
Estimated Suspension System Frequencies 16
Handling Procedures . 17
STRESS ANALYSIS 18
Discussion of Ground Rules 18
Discussion of Load Factors 20
Model Loading Conditions . 2 1
Flow Chart for Model Handling 22
Basic Model Configuration and Geometry 23
Model Interface Load 25
Orbiter Mass Distribution . 27
Orbiter Handling Conditions 28
Orbiter Section Properties 29
Orbiter Loads 30
Orbiter-to-Tank Interface Loads 34
SRM-to-Tank Interface Loads 36
SRM Ring Loads 42
SRM Drag Strut Analyses 43
APPENDIX A - Loads Due to 1 Lb. Oscillating Force at Engine A-1
i iAPPENDIX B - Resonant Frequency of Body Suspended From Two Angled Wires A~2
iii
FIGURES and TABLES
TABLE I
TABLE II
FIGURE 1
FIGURE 2
FIGURE 3
FIGURE U
FIGURE 5
FIGURE 6
FIGURE 7
FIGURE 8
FIGURE 9
FIGURE 10
FIGURE 11
FIGURE 12
Weight Statement for Configuration 619
Drawings of 1/8 Scale Model
Mated Flight System Design 619
External Tank Inboard Profile
Comparison of Model and Prototype External
Tank Area Moments of Inertia
Comparison of Model and Prototype External Tank
Cross Sectional Areas
Approximate Prototype LCU Tank Dimensions
SRM Inboard Profile
Orbiter Structural Arrangement '
Comparison of Model and Prototype Orbiter
Cross Sectional Areas
Comparison of Model and Prototype Orbiter
Moments of Inertias*
Assembly Drawing of 1/8 Scale Structural
Dynamics Model
Schematic of Model Suspension and Leveling
Systems
Shuttle Model Suspended
PAGE
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iv
MODEL
SUMMARY
The basic objectives of the l/8th scale Preliminary Structural
Dynamics model are to:
o Provide early verification of analytical modeling
procedures on a Shuttle-like structure.
o Demonstrate important vehicle dynamic characteristics of
a typical Shuttle design
o Disclose any previously unanticipated dynamics
problems
o Demonstrate optimum configuration changes for
eliminating critical problem areas
o Provide for development and demonstration of cost-
effective and efficient prototype testing procedures
The design objective for this task was to represent important
structural dynamic characteristics (discussed in CR 112196) while
keeping the fabrication costs low. The basis for the model was
the Grumman proposed Design 619 Space Shuttle which was a 4.8M Ib.
GLOW 182 ft. long parallel burn configuration. Simplifications
included extensive use of constant thickness unstiffened skins
in place of variable thickness skin-stringer-frame construction,
frames designed as back-to-back channel with elements formed
from cut flat plates fastened between them to act as fittings
in place of machined frames, and simple tubular struts with
standard AN end fittings for interstage members in place of more
elaborately formed members. These and other simplifications in
the design resulted in locally stiffer and heavier areas than
would occur in a full replica model, however, they were necessary
to keep the fabrication costs within target. Structural joints
in the model were designed to be simpler and stiffer than the
prototype in an effort to avoid the extra flexibility which has
occurred in replica scaling down to thin gages.
The model consists of k elements. An Orbiter, and External
Tank and two Solid Rocket Motors. Since investigation of
hydroelastlc effects was an important objective in the test program,
and since the ET flexibility and weight was expected to be the most
significant factor in the low frequency modes, the ET was the first
element to be designed. !
The ET is 39.5" in diameter and 237.8" long and consists of
a forward LCL tank, an intertank skirt with provisions for
attaching the.SRMs, an aft LH_ tank with internal frames to
support the Orbiter attachments, and an aft skirt providing the
aft SRM attachment. In order to avoid buckling of the LIL tank
during vertical suspension of the model, it was designed to be
supported from the intertank skirt. The L0p tank is of welded
monocoque constant thickness design 78.3" long. The forward
portion is conical while the aft is cylindrical. Both regions
have .02" thick walls. The bottom consists of an .016" thick
lower dome. The intertank skirt is 29" long cylinder of
.050" wall thickness. There are 3 frames to accept the SRM and
the model suspension system attachments. The LHp tank is a iW
long cylinder with 10 frames, 3 of which take the orbiteri
attachments and the remainder required to prevent buckling. The
skin thickness is .025" in the upper regions of higher load and
.016" for the remainder. The aft skirt is a simple .020" thick
cylinder 13" long with a single frame to take the SHI attachments.
The SRM consists of a forward skirt with ET attachment
provisions, a propellant cylinder, and an aft skirt with both ET
attachment and hold-down support provisions. The propellant
cylinders are 147.3" long, 19.5" in diameter with 0.2" thick wans.
The propellant simulation proposed is inert PBAN which has an inert
salt in place of the oxidizer used in the active material. Six
cylinders are proposed, two each to represent full, maximum qO< ,t
and burnout conditions. The forward skirt is a 23.6" long .05"
thick cylinder with a forward and aft frame to provide for ET
attachment. The aft skirt .'consists of a cylindrical and conical section. The
5" long by .062" thick cylindrical section contains the SRB to
ET attachment frame. The 22" long .062" thick conical section
progresses from 19-?" diameter to 30.2". Four tapered longerons
attached to this cone represent the hold-down supports.
When the ET and SRM designs were almost completed,
manufacturing cost estimates were made to determine the fabrication
cost target for the Orbiter. The Orbiter representation which
met this requirement consisted of a forward non-circular shell
representing the cabin, a midsection payload and wing area, and
an aft section providing a representative engine and fin support
structure. The fuselage external lines are simplified to •• '
minimize curved sections, the sides and bottom are flat as are
the surfaces of the wings and fin stub. The forward .020" thick
shell is 17.25" long and varies from 19" to 26.4" deep. Longerons
are extended along the sides and slightly forward to constitute
the forward hoist point when lifting the Orbiter separately from
the remainder of the model. The midsection is 102.5" long and
consists of a series of U shaped frames spaced every 10" running
up to shoulder longerons along each side. These frames are covered
by .020" thick skin. Payload attachment provisions are made at 4
locations along the length. The payload door is a semi-cylinder
.016" thick divided into 7 segments along its length by V shaped
angles which permit the door to carry torsion but not bending.
The wings consist of 6 spars 2.5" deep at the tip and 6" deep
at the root extending 1*9" from the fuselage and covered by .02"
Th': uit oc-ctiori of the orb It or IB 20" long and is designed to
represent the basic load carrying structure between the orbiter engines
and the ET thrust load attachment. At the aft end is an open frame
which supports the aft fin beam. Forward of this is the engine bulk-
head containing provisions for mounting representations of one upper
and two lower engines. A sloping deck between the engine bulkhead and'
the aft payload bulkhead provides a load path from the upper engine to
the shoulder longerons. Two struts provide a path of proper stiffness
between the lower engines and the ET thrust attachment. Stability of
the sloping deck and the strut attachment points is provided by vertical
channels attached to the engine bulkhead.
MODEL DESCRIPTION
Basically the design is a simplified l/8th scale model of a
parallel-burn Shuttle, which full scale is represented by the
Grumman Design 619 (or Gill) shown in Figure 1. A summary, weight
statement for this prototype is listed in Table 1.
In simplifying the design, a major objective was to keep the
fabrication costs within target while retaining as many of the
significant structural characteristics important for dynamics as ipossible. The guide lines used are outlined in Section 1 of
CR112196 prepared under HAS 1-10635- .
External Tank
The prototype design used as the basis for the model is shown
in Figure 2. The principal parts of the external tank are the
oxygen (LC>2) tank, the intertank skirt, the hydrogen (LHp) tank,
and the aft skirt. '
Initially, in designing the model external tank, an effort
was made to adhere to the scaled down cross-sectional area and
the moment of inertia in bending for both the major structural
elements and the interconnecting members. This proved to be
difficult in the LIL tank when an effort was made to provide
for a test condition where the configurations without the SRM's
were supported "by the orbiter engine base. Buckling occurred
if scaled down thicknesses were used even with the addition of
rings at less than 2" spacing. Suspending the model from the
intertank skirt at the interstage connection and thereby putting
the LHp tank in tension eliminated this problem below the
suspension point but buckling was still present above the
suspension point in the intertank skirt. Therefore, after
discussion and a NASA review, the change of support point to the
intertank skirt was made and the intertank skirt thickness was
increased in gage until, buckling was no longer a problem. This
approach also minimizes fabrication costs. .The comparison in
areas and inertias between the prototype direct scaling and the
model design is shown on Figures 3 and k. It was felt that the \
resulting configuration would still provide a good check of
analytical methods, and that the major structural dynamic
interactions between the model SRM, external tank, and orbiter
would still demonstrate the type of behavior anticipated on the
prototype.
The prototype LCU tank was monocoque construction with variable
tapered skin thicknesses as shown on Figure 5. The model tank is
also monocoque. The lower dome is formed of 3 spherical segments
to save the tooling required for an elliptical shape. The
thickness which varied in the prototype, is kept constant in the
model at .016". A thicker section is retained adjacent to the Y
ring to permit welding. The model Y ring is considerably larger
than the scaled down prototype dimensions and is made in simple <
tapered shape to expedite analysis. The cylindrical section of
the model LCU tank which if scaled directly, would require
variable-thickness (.023 to .016") is kept constant at .020". The
conical section (scaled prototype design .016" to .012") is kept
constant at .020" to avoid welding difficulties. The upper dome
which, if scaled directly, would have been .0065" is .025" in the
model in order to limit welding and handling difficulties. The
upper dome is a single spherical section. A removable upper
cover is added to permit inspection and cleaning of the model. The
dimensions of the L0? tank are adjusted to provide proper scaled
L02 weight using HO.
The Interbank Skirt in the prototype was a ring-frame stiffened
cylinder with one very large frame stiffened by internal struts
(Section B-B, Fig. 2) to carry the SRM symmetric lateral (Y
direction) load. These strut cross-sectional areas would be .16
to .19 sq. in. if directly scaled, but in order to achieve
commonality of members throughout the model they are increased
to .26 sq. in. Back-to-back channels are used in place of tubes.
The skin gages required in the prototype varied from a minimum
of .070" to a maximum of .2". Scaling these values down to .009"
to .025" for the model would result in buckling in this area
under handling and vibration test induced loads. Therefore, to
avoid the complexity of many rings at close spacing, and chem
milling to various thicknesses, it was decided to select a
..050" aluminum skin which is the minimum for a uniform thickness.
There are, therefore, only 3 frames required in this area of the
model. The SRM axial loads are applied to two fittings on each
side. One of these fittings is shown in detail in Section B-B i
on Figure 2. la the model, this fitting is also used to suspend
the configuration. Therefore, the model is heavier than a scaled-
down version of the prototype would be. In simplifying the
design to reduce machining costs, still more weight and stiffness
is unavoidably added.
The LH tank prototype, as shown in Figure 2, was a ring-frame
stiffened cylinder with 3 major frames and fittings to accept the
orbiter induced loads. The skin thickness if scaled directly,
would vary from .026" to .015". To simplify construction of the
model, the skin is either left at. .025" or chem milled to .016".
The total number of ring frames and ring stiffeners is limited to
10, about half the number in the prototype. This is feasible
because the buckling loads were low. All three ring frames are
back-to-back channels having the same channel section in order to
save tooling costs. The I (area moment of inertia) selected for
each of the frames is 0. 5 in. which is approximately representa-
tive of the scaled prototype value of 0.1*9 in. for the forward
interstage frame. The end domes in the model are .020" in place
of the .009" scaled from the prototype in order to avoid welding
and handling problems. They are formed with the same tooling as
the LOp tank dome and, therefore, unlike the prototype,- they
have the same geometry. The internal struts in the frames '
which distribute the aft orbiter loads are of the same geometry
as the prototype but made from back-to-back sections in place
of tubes in order to save tooling costs. The cross-sectional area '
of the model internal struts is larger than the scaled prototype2 2
value (.26 in. in place of .23 or .19 in. ). The prototype
orbiter drag fitting was as shown in Detail P of Figure 2. In
the model, the drag fitting has the same effective line of action
but is made of a single machined straight sided element which is
•stiffer and weighs more than the scaled prototype value.
The aft skirt in the model is simpler than the prototype
since there is no beading of panels or ventral fins and only
one ring frame which distributes the aft SRM loads. This is2
stiffened by lateral struts which would be .19. to .22 in. if2
scaled from the prototype, but which are .26 in. in the model
to save the cost of additional tooling. There is no full
bulkhead installed in the model as there was in the prototype
since the bulkhead was not considered significant for the structural
.dynamic characteristics.
8
SRM
The prototype design used as the basis for the model is shown
in Figure 6. The central cylindrical section which in the
prototype is made of 6 segments of .57" thickness steel is modeled
by a single 0.2" thick walled aluminum cylinder. The propellent
is modeled by an inert propellant in which the oxidizer replaced
by an inert salt. No structure was considered necessary to
represent the upper dome of the prototype. The lower dome is
represented by a conical section for simplicity.
The forward skirt in the prototype SRM was designed as an
orthogonally stiffened steel cylinder with closely spaaed longerons,
and with 5 ring frames spaced about 26" apart. Skin thickness
varied from .2" to .06". In the model, this is represented by
an unstiffened .050" thick uniform aluminum cylinder which
represents average thickness, and because this thickness will
prevent buckling under any of the model loading conditions.
The forward skirt of the model has frames consisting of back-to-
back channels located at the forward and aft intersections of the
interstage struts with the skirt in order to take the SRM-to-
external-tank loads. In the prototype, the forward frame had an2 2
area which varied from about U in. to 10 in. , with a correspondingU , hvariation in I from 125 in. to 490 in. . The aft frame area
2 2 Uvaried from 1.8 in. to 6.2 in. , and the I from 18 in. to
ji106 in. . On the model, both are represented by the same frame with
2 kan area of .27 in. and an inertia of .28 in. which, when scaled2
up to prototype size and material, would represent a 5.7 in. areaIt
and a 30 in. inertia. This is considered within the proper
range for the forward frame.
The prototype geometry of the interstage strut fittings is
retained but the model fittings are simplified. The interstage
strut attachment lugs are fabricated from plate to save machining
costs. This makes the model elements heavier than scaled down
prototype design. The SRM recovery system is represented by
lumped weight attached to the forward ring.
The aft skirt of the prototype which contains the hold-down
fittings and SRM-to-external-tank fittings, was designed as an
orthogonally stiffened conical frustrum with stringers every h degrees
and ring frames approximately every 2k". The prototype material
was steel. Skin thicknesses varies from 0.3" to .075". This
structure is simulated in the model by a conical .062" unstiffened
uniform aluminum skin. The prototype had U major tapered longerons
extending from the hold-down fittings to the SRM cylinder each with2 14-a maximum area of 10 in. and an inertia of 389 in. . These are
represented in the model by back-to-back channels which when scaled up2
to prototype material and dimensions have an area of 8.6 in. and
an inertia of 280 in. . Above the conical section of the aft SRM
skirt, the prototype had a short cylindrical section containing
the fittings for the struts linking the SRM to the external tank
and containing the ring which fastens to the aft portion of the
SRM propellant cylinder. On the prototype, this ring was a double
U shaped section about 10-|-" deep by k- " wide with an area of2 kabout 20 in. and an inertia of about 290 in. . A simpler single U
2section is used in the model having an area equivalent to 13 in.
Ij.and an inertia of 95 in. when scaled to prototype dimensions and
material. The short cylindrical section of the model forward of
the U ring has the same skin thickness (.062") as the conical
section, and is terminated in a ring for attaching to the model
SRM cylinder. This area, while not representative of the
10
prototype, is designed to have stiffness compatible with the
upper portion of the conical skirt.
Orbiter
The prototype structural arrangement used as a basis for the
orbiter model is shown on Figure 7.
The aft portion of the model fuselage has similar geometry
as the prototype but is designed with only two full bulkheads,
one aft frame and two intermediate frames in place of the larger
number on the prototype. The upper cutouts in the prototype
which providedvfor the QMS and abort SRM's are not modeled,
instead full frames are used. The external lines of the model are
simplified and straightened compared to the prototype to reduce
fabrication costs. The cross-sectional areas of the struts between
the lower engines and the aft interstage fitting are scaled
directly from the prototype as were the side longeron areas. A
simple fin structure which extends up to the location of the fin
e.g. is also included in the model. The skin gages in the
bulkheads (.032" aft and .OhO" forward) are adequate for
fabricating, thick enough to avoid buckling, and heavy enough
to simulate some of the non-structural weight in the prototype.
The in-plane areas of the prototype bulkheads are not scaled
for .the model in order to avoid structural complexities and because
it is not considered significant in establishing the primary
structural dynamic characteristics of the model. All side skins
of the model fuselage are .020" thick, which is adequate to
avoid buckling. The scaled down prototype dimension for the
fuselage side skins is .012" but this would require intermediate
rings and longerons in the model and increase the cost of the
fabrication. The deck of the model which distributes the upper
engine loads to the longerons is .016" thick and correctly scales
the prototype stiffness (area/length).
11
The prototype fuselage mid-section consisted of closely
spaced frames covered "by corrugated outer skin carrying TPS tiles.
The model has similar but simplified geometry and structural
arrangement. The model consists of series of more widely spread
U shaped frames spaced every 10" covered, by a .020" side skin
and .025" bottom skin. This skin thickness is enough to prevent
buckling under static load but it results in a larger area and
inertia than the scaled prototype values as shown in Figures 8
and 9- Payload support provisions are made at h different stations
to permit variations to be tested. The longerons of the model are
designed to furnish the proper scaled area for the aft end of the
orbiter. It is kept at a constant area for the entire length of
the fuselage to limit fabrication costs. In order to simulate the
prototype weight, about 0.8 Ibs/inch, including the structure,
•would, be required on the model. This is accomplished by increasing
the thickness of the frame webs where stiffness is not significantly
affected.
The wing of the model consists of 6 beams covered top and
bottom by flat plates. Wing root connections are made by bolts in
shear through the webs of the beams and machine screws connecting
the top and bottom wing skins to the fuselage. This method of
attachment simulates the prototype. The prototype had a corrugated
double skin with an average thickness of .16" to iV. This
structure is simulated in the model by a .020" sheet. Since the
wing depth of the model is properly scaled, the prototype inertia
at the wing root is properly represented on the model. A constant
skin thickness is used over the entire model wing to control costs
and this does give the proper order of magnitude for inertia since
the beam depth decreases toward the model wing tip. The proper
weight for the wing including the TPS panels is simulated on the
model by adding thickness to the webs of the spars. The model
12
wing does not have chordwise trusses or beams (typical of the
prototype) so that loads applied away from the outer periphery
are not distributed properly between spars, but loads applied
i at the outer edges are properly transmitted by the top and bottom
covers. Therefore, realistic modes are anticipated if shakers
are kept at the periphery of the wing. The RCS wing tip pods
of the prototype are simulated by lumped weights on the model.
The orbiter interstage fittings duplicate the prototype
geometry using simplified components.
The orbr;er model payload bay door consists of a removable
7 segmented semi-cylindrical cover skin which can take loads in
torsion but not in tension and compression, thereby simulating
the structural properties of the prototype door in. a closed and
locked position. The minimum skin gage which could be used for
the door in order to prevent buckling is ..016", which is quite
thick compared to the scaled prototype value of .00325".
The forward fuselage in the model consists of a tapered
non-eiroular stiffened shell extending through the cabin location.
The two side longerons of the fuselage mid-section are extended
through this area in order to provide a forward orbiter model
hoist point. There are provisions for attaching weights simulating
the forward equipment. Additional stabilizing stiffeners are
added to the model to prevent buckling. Local stiffnesses are not
considered significant for vehicle structural dynamic
characteristics and are, therefore, not scaled.
Complete design drawings of the l/8th scale dynamic model are
available f-t the Dynamics Loads Branch, Loads Division, NASA/Langley
Research Center. An assembly drawing of the model showing the major
components and overall dimensions is presented in Figure 10.
13
SUSPENSION SYSTEM FOR MODEL
Suspension System For Model
The objectives of this system are to provide support which does
not interfere with measurements of the unrestrained structural dynamic
characters tics, such as mode shapes modal frequencies ;and damping
during tests, while keeping the support induced buckling and handling
loads to a minimum. The system should also .permit the ready disassembly
and changes in the model elements required by the test program. The
system must be compatible with the current support structure (Backstop)
in the NASA Langley Dynamics Research Laboratory, and should be relatively
inexpensive to implement.
Suspension Concepts
Various support concepts were considered for use on the model.
These included the systems analyzed by R. W. Herr and H. D. Garden in
USAF RTD-TDR-6"3-]H9T (Sept. 1963). Vertical suspension of the launch
configuration is necessary to obtain the proper hydroelastic interactions
in the L0p tank* Successful use of air springs for large vehicles at
Grumman prompted the adoption of 0.5 bz units, remaining from the IM
program to provide vertical isolation. Horizontal isolation is compli-
cated because unlike previous axisymmetric models, the Shuttle center
of gravity shifts laterally with reduction in fuel so that the attitude
changes for most practical vertical suspensions. Base support springs
as used on the full scale Saturn V Dynamic Test Vehicle were considered
too expensive and a cable system was therefore adopted. A single
cable was impractical because no adequately strong suspension point was
available. Therefore the system described below was adopted and
designed.
Recommended Suspension Scheme'
A modified two cable suspension system shown schematically in
Figure XL was selected. The modification consists of a bridle positioned
between each suspension cable and the model. The bridle is routed
thru a sheave on the suspension cable and each end is attached to an
SRM interstage fitting on the same side of the ED taak. This arrange-
ment offers the advantage of supporting the model at the same four
points at which the SRM thrust is introduced on the full size vehicle,
mini,mi.zing the effect of the suspension loads on the model and permit*--
ing the model to assume its eqilibrium attitude at any level of pro-
pellant loading.
Within the restraints shown on Dwg AD 383-500* "the primary sus-
pension system may "be routed in any number of ways thru a system of
sheaves mounted on the upper backstop structure. One such routing is
presented on Figure 11 .
Provision for changing the SRM propellant level between tests
was an important consideration in the design of the. overall system.
Discussion with KASA concluded that changing propellant cylinders,
(relatively long and heavy masses,)on both SRM's with the model sus-
pended was not desirable, and, that the approach should be to first
lower the model and to support it vertically on the floor prior to
replacing the cylinders. Since the suspended model can be in any of
three possible attitudes this requires that it be oriented vertically
prior to lowering." Also shown in Figure 11 is a proposed
routing of the model leveling • cable system. It should be noted that
this routing is proposed to run parallel to that of the suspension
system in order to maintain the attitude attained by use of the
leveling system while the model is being raised or lowered.
Both cable systems are interconnected by a hydraulic ram and
sheave arrangement, also shown in Figure 11 . The hydraulic ram is
proposed, permitting remote operation and selection of rate, however,
any mechanical device of adequate capacity, such as a chain hoist,
could be used instead . The ram changes the model orientation by
varyiiig the distance between the two sheaves, one being held fixed by
the cable to the actuator on the floor,by a winch or ram or any other
satisfactory device, the other connected to the orbiter and free to
move. During tests the leveling ram is fully extended permitting the
model to assume its equalibrium attitude. The slack leveling cables
may be left attached to the orbiter or disconnected. The model may be
15
raised or lowered in any attitude by means of the floor mounted actuator.
Drawing AD 383-500-1 showing the extreme positions anticipated
for the range of weights to be tested with and without SRM is shown
on Figure 12. Also shown is the clearance, anticipated. The orbiter
alone could be readily suspended from the forward attachment point,
and an aft handling point at the top of the fin is available for
support during any required movement.
Estimated Suspension System Frequencies
The fundamental model free resonant frequency is expected to be
about 8 hz. Therefore suspension frequencies below .8 hz would be
desirable for the support system. The IM airsprings will furnish
a .5 hz vertical suspension which should be adequate. Laterally the
combination of rocking and pendulum motion- is anticipated. To
estimate the lateral frequencies, the model was analyzed in two planes
separately assuming the vertical airspring was not effective.
In the plane where the vertically suspended model is viewed from,
directly in line with left wing, the right hand bridle attachments
are directly behind those on the left hand side and the system acts
like a simple physical pendulum with the center of gravity below the
point of support,- as described on page 250 of "Advanced Dynamics"
by Timoshenko and Young. The two resonant frequencies were calculated
for 5 weight conditions from full-up (9 32 Ibs.) to just prior to
decoupling the. external tank (675 Ibs). The resonant frequencies for
the lightest condition were 0.6l hz and O.lB hz, while those for the
heaviest were 0.37 hz and 0.19 hz.
In the plane at right angles to this, where the suspended vehicle
is viewed from directly in line with the orbiter fin no adequate
expression for the resonant frequencies was found. An approximate
expression was developed using Lagrange's equations and assuming small
motions and equivalence of small angles as shown in Apprendix B. The
resonant frequency in this plane varied from .20 to .26 hz for the range
of weights from full (9 32 Ibs.) to almost empty (6751bs). A more
detailed analysis would include the effects of the vertical flexibility
16
however the resonant frequencies for the idealized cases are considered
sufficiently low to provide confidence that the isolation furnished by the
suspension should be adequate.
Handling Procedures
The procedure recommended for the initial assembly of the model is listed
on Figure 10. The weights and angles anticipated for various loading conditions
are shown schematically ia the Stress Report on page 22 "Flow Chart For Model
Handling".
17
AMW
PAGE 18
STRESS ANALYSIS
The Stress Analysis includes the following:
1. A discussion of ground rules concerning the handling of the model.
2. Presentation of th.e loads associated with these ground rules.
3. Documentation of model geometry at internal interfaces, and the determination
of model loads for which the model structure was checked.
U. Analysis of a typical SRM drag strut.
The detailed analysis of the "basic internal model is not presented herein. In
designing a 1/8 scale model, true scaling results in margins of safety up to 8
times those on the prototype for the same accelerations. For this model, for
cost control and design simplicity, many critical structural areas show even
greater margins of sai'ety. Each drawing was reviewed to assure adequate
factors of safety and stability for all defined load conditions. It is not
considered necessary Co include all detailed calculations in this document.
Discussions of Ground Rules . . . '.
Handling conditions are presented on Page 20.
All raising and lowering of the model shall be accomplished with the Op
(water) tank of the external tank drained and empty; and with the model in a
vertical orientation with the orbiter supported vertically from the nose
fittings provided at Orbiter X Station 1*6. These stipulations are necessary
to (l) prevent buckling of the intertank skirt between the pickup fittings and .
the 02 tank Y ring and (2) to prevent compression bucking in the lower skin
of the orbiter.
Loads Induced by Handling
The load factors acting on the dynamic model are tabulated on Page 20.
These include mode survoy as well as model ground handling conditions.
Flight configurations are designated by symbols "A" thru "G" as given on
Page 21. Condition "A" is representative of the prelaunch configuration.
Flight conditions progress until the last flight configuration condition "G" is
achieved. The appropriate model weight for each flight condition is also
listed on Page 21.
A "Flow Chart for Model Handling" is presented on Page 22. The chart lists
the configuration sequences that the model sees in the test program. Configured
tion "A" may proceed to configuration "F" by means of 30 steps. Each step is
designated by the step number encircled and an adjacent arrow. All 1 g support
loads and the inertia loads on the model components for each step are given on
Page 26. These support loads occur either at the base of the SRM or at interface
G A C 3 2 8 A R E V 2 REPORT WAS 1-10635-11
OATE ^ October 1972
CODE 26912
PACE 19
points 1 & U. The interface points are located on Page 2k. Ultimate loads for
design for these locations are shown on Page 26.
Orbiter Loads
The orbiter weight distribution & inertias are presented on Page 27 & 29.
The loads induced oh the orbiter during handling are presented on Page 28. Loads
induced at the interface locations due to a unit load applied at both the center
of gravity & the forward nose hoist location are shown. The shear axial load &
bending moments for a 1 g axial and lateral applied force are then determined as
shown on Pages 30 to 33- These form the basis to check the strength of the
orbiter fuselage.
Orbiter to external tank interface loads are obtained by applying unit loads
at the interface locations and calculating the strut member loads & the tank
support point reaction forces as shown on Page 3U. The loads at these interface
locations due to critical handling conditions on the orbiter are obtained by
.multiplying the appropriate values on Page 26 by the unit factors on Page 28 &
are listed on Page 35- These are then converted to the member & fitting loads in
the lower table on Page 35.
SRM Loads
A similar procedure is followed in determining the SRM induced loads on the
external tank interfaces. The geometry & distribution of forces to the inter-
faces is shown on Page 36. The loads at the tank reaction points & in the truss
members for a unit axial (l Ib.) load applied at the point designated 5' is shown
on Page 37. A similar distribution for lateral (Y & Z) loads applied at the SRM
center & for lateral loads and moments applied at the center of the external (HD)
tank are shown on Page 38 for the forward interface & Page 39 for the aft inter-
face. The reaction forces have been translated into axial, radial, & tangential
forces & strut loads & summarized on Page UO. The critical handling conditions
from Page 26 are then multiplied by these factors to give the ultimate loads as
listed on Page Ul. The primarily axial handling loads are combined with the
lateral handling loads to give the design values listed on Page Ul. A similar
calculation for the loads in the SRM at these interfaces gives the values
shown on Page U2.
Analysis of a Typical SRM Drag Strut
A typical analysis is shown on Page 1*3. The loads applied are determined
on Page Ul-. Two cross sectional areas are checked and the margin of safety
is .27.
O A C 3 2 6 A R E V 2 REPORT MAS 1-10635-11
f t"TI '"M DATE 3 October 1972
CODE 26512
PAGE 20
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TABLES AND FIGURES
TABLE I
WEIGHT STATEMENT PAGE 1 OF 1
CONFIGURATION ' 6l9
CODE
1 ' "
23U56.78910111213lU15161718
2021222325
T
2125222526272829
iSYSTEM ; STACK ELEMENT
ORB."WING GROUP " 18017
! TAIL GROUP 3186: BODY GROUP , 38lH6' INDUCED ENVIR. PROTECTION . 291 5
LANDING, RECOVERY, DOCKING i 96l3PROPULSION-ASCENT 2l8l6PROPULSION-CRUISE 217PROPULSION-AUXILIARY 5738PRIME POWER 367UELECT CONVER. & DISTR. 2975
; HYDRA CONVER. & DISTR. 15 6SURFACE CONTROLS . 1538
1 AVIONICS ' 6615: ENVIRONMENTAL CONTROL UUl8: PERSONNEL PROVISIONS 1068: RANGE SAFETY & ABORT
BALLASTGROWTH /UNCERTAINTY 13068NOSE CONESEPARATION & DEORBIT 100
SUBTOTAL (DRY WEIGHT) 160880
PERSONNEL . ' lU20CARGO (LANDED) hooooORDNANCE (IGNITER)RESIDUAL FLUIDS 1803RESERVE FLUIDS ' ( LANDED ) 897
SUBTOTAL (LANDED WEIGHT) 205000
A CARGO UP 25000FINS /SKIRT (STAGED)RECOVERY SYS. (EXPENDED)IGNITER (EXPENDED)RESERVE FLUIDSIN FLIGHT LOSSES 6 32PROPELLANT-ASCENTPROPELLANT-CRUISEPROPELLANT-MANEUV . /ACS 27 396SEP. SYS. /NOSE CONEINSULATION /LINER
ELEMENT GROSS WEIGHT 263828ABORT SRMGLOW ,
_ _ _ _ _ SRM TANK
290312 501995999
586027900 2350 :
:
i1720 . . U69 i
i
500 i!
lUS^O 1300113U
1200 ; 1*85
3 1832 i 66305 i~~ ~ "" ^-— — — " I ~~ ~~ ^^^ — L~ ~
j :1 i
ti ii
1028I 8186 •
7333I
1
3U2860 , N. A.
1
i 1039621220
900 t :
'
2 52932 1560052I
i • :
5600 I ;21200
28U712 165227268500
1*829312 i
TABLE II
DRAWINGS OF 1/M fiC'AI.F: MODEL
DRAWING NUMBER
AD383-500
-501
-502
-503
-50U
-505
-506
-507
-508
-510
-511
-512
-515
-516
-517
-518
-520
-521
-522
-525
-526
-527
-528
-529
-530
-531
-532
-533
DESCRIPTION
Model Assembly Suspended (3 Sheets)
Shuttle Model Assembly
External Tank Assembly
SRM Assembly
Orbiter Assembly
L02 Tank Assembly (2 sheets)
Intertank Skirt Assembly
LHg Tank Assembly (2 sheets)
Aft Skirt Assembly
SRM Forward Skirt Assembly
SRM Propellant Cylinder Assembly
SRM Aft Skirt Assembly
LH2 Tank Fitting Installation
Rings for External Tank
Intertank Skirt Frame Assembly
LH2 Tank Frame Assembly
External Tank Aft Skirt Frame Assembly
SRM Rings
SRM-to-External Tank Thrust Fittings
External Tank-to-Orbiter Thrust Fitting
Orbiter Forward Section Assembly & Installa-
tion
Orbiter Payload Bay Cover Assembly & Installa
tion
Orbiter Payload Module Installation
Orbiter Aft Section Assembly
Orbiter Wing Installation
Orbiter Fuselage Side & Bottom Skin Panel
Assembly and Installation
Orbiter Keel Assembly and Installation
Orbiter Wing Beam Carry-Through Assembly
Orbiter Aft Interstage Fitting Assembly
TABLE II (continued)
DRAWING NUMBER
-53k
-535
-536
-53T
-538
-539
DESCRIPTION
Orbiter Engine Support Bulkhead Assembly
(2 sheets)
Orbiter Fin Stub Installation
Orbiter Fuselage Forward Frame Assembly
Orbiter Abort SRM Installation
Model Cosmetic Lines (2 sheets)
Orbiter Engine Bulkhead (Station 180.009)
Fittings
Note:
1. Two (or more) copies of each of the above drawings have been
submitted separately to NASA/Langley for review.
2. These drawings are available from the Dynamic Loads Branch,
Loads Division, NASA/Langley Research Center, Hampton, Virgina 2336^
EXTERNAL TANK ORBITERORBITER BURNOUT C.G. INSERTION
(505SECI C.G.
\ >
^7v V \ \\. x ' S V N l\ XX
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3-555
Fig. i Mated Flight System
SEAL FWDSRM rn -|MFATTACHMENT (4) G02 LINE
L02LINEX
ELECCONDUIT
SECTION L-L
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VENT RELIEFVALVE
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G02
PRESSLINE
VENT VALVE SECTION B-B x
ACTUATIONGHeTANK
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DETAIL P
rB LO2 MAIN c
'NOSE CAPSEPARATION PLANE
NOSE CONE- 220.5
-RETRO MOTOR
PYRO INITIATOR""DRIVE (2)
PY'RO BATTERIES (2) | FEED LINE GH2 VENT I r FWD ORBITERDEORBIT ELECTRONICS (2) 232.4 I / ,PR|SS l!/ SUPT FITT NG(SEQUENCER. POWER BATTERIES, f^-|NTER TANK- '^ /PRESS ^/ SUPT I-1 III Nfa
LH2 TANK ACCESS ROUTE
EXPLOSIVE BOLT SRM STRUT
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^——•LH2MAIN /-ANTI-VORTEX BAFFLE ASSY
SEE DETAILG
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SEE DETAIL J
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3-608
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AFT SKIRT INSTL. SEPARATION SYSTEM
DETAIL K
DUAL LINEARSHAPED CHARGE
PYRO LINE (2)PYRO INITIATOR (2)
Fig. 2 External Tank Inboard Profile
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APPROXIMATE PROTOTYPE L02 TANK DIMENSIONS
Figure 5
51
IT.T. PORT
k
AFT STAGINGROCKET MOTOR PACKAGE
., PAD SUPPORTJf & HOLD-DOWN FITTING
r 1(tPAD
<t DROGUEPILOT MORTATTf
MAIN CHUTE FITTING(3) (SEE SECT A-A FOLOCATION)
INSULATIONASBESTOS SILICAFILLED NBR
RINGS / r-.571TANKOML(REF)
NUTBOLT
EXPLOSIVE INSERT
CENTERING,WASHER
j FWD. THRUST MOUNT':' / INTERFACE PLANE FORWARD THRUST
FITTINGLUBRICATEDSURFACE DETAIL
SEPARATION BOLT
<l AFT MOUNT FRAME
PBANPROPELLANT3-°-\Z.NC
CH ROM ATEPUTTY
AFT SKIRT SPLICE(PINNED CLEVIS JOINTa RETAINER STRAP)
LINER (.065 NOMCOMMAND CONTROL SPLICE PINANTENNA (TYP) (2) RETAINER BAND(SEE SECT. A-A FOR LOC.)
RECOVERY SECTIONEPARATION
POINTCHUTE RELEASE -EXPLOS. BOLT
THROAT SECTION
69
MORTARMAIN CHUTE PILOT
MAIN CHUTE STOWAGE(3 CANISTERS) (1 CHUTEPER DEPLOYMENT BAG/CANISTER)
40 ^-PINNED CLEVIS SPLICE
SEGMENT260(4 SEGMENTS)
150FWD CLOSURE SEGMENT
147AFT CLOSURE SEGMENT
PINNED CLEVIS SPLICE FWD. THRUST SKIRT
SHAPED CHARGE (TYP)
DROGUESWIVEL FITTING
SHAPED CHARGE (TYPlPIN JOINTW. RETAINER RING
NOSE FAIRING
EQUIPMENT BAY(TYP) (2)(SEE SECT. A-A FOR LOC)
ELECTRICAL POWER'& RECOVERY SYST.
TRACKING (C-BAND) ANTENNA (4)-*.
F —INSTRUMENTATION& FLIGHT SAFETY
DROGUESWIVEL FITTING
EQUIPMENT BAYINSTRUMENTATION& FLIGHT SAFETY
COMM/CONTROLANTENNA
TANK HORIZONTAL q.
76
RECOVERY BEACON'ANTENNA (2) (STOWED)
9. MAIN CHUTEFITTING
MORTARMAIN PILOT
EQUIPMENT BAYELECTR POWER& RECOVERY SYST
FWD THRUST FITTING<i MAIN CHUTE (FIXED TRI-POD)FITTING
TRACKING (C-BAND),NTENNA (4) 90° SPACING
PROV FOR ANTENNA INSTRH. SRM
' -*— VERTICAL (Z)(R.H.SRM)
DROGUE\ > STOWAGE MORTAR
— DROGUE PILOTS
\ |.;U. ,_,^ q^STAGINGSRM'S
/ '•() C MORTAR^ATTIT. SENSOR
BALLUTESI3)DIRECTION-
XA. SENSOR (3)
* ^COMM/CONTROL ANTENNA(INSTALL ON VERT. C )
<t VERTICAL (Z)(LEFT-HAND SRM)
FWD THRUST FITTING(SWING-AWAY BI-PODNEAR-ORBITERSIDE)
<L SRM ATTACH GEOMETRY
-f 76BREAK-AWAYCONNECTOR
,(5LLH. 4 RH) / IT* AFT MOUNT' I 7r' LOWER I
1 AFT T.T. PORT
r 76-76-
AFT ELECTR. COMPARTMT(SQUIB FIRE CIRCUITS)
AFT STAGINGMOTOR PACKAGE
<t HORIZONTAL
PAD SUPPORT& HOLD-DOWN FITTING
r3^536
MAIN CHUTE C MAIN CHUTECANISTER (TYP) FITTING
RECOVERY SECTIOND-D
FORWARD MOUNT
E-E
AFT MOUNT f-fAFT SKIRT
Fi SRM Inboard Profile52;
Structural Weight Summary
Component
FuselageNoseCrew CompartmentMid FuselageSide PanelsPayload DoorsFwd Wing Carry-ThruAft Wing Carry-ThruAft FuselageQMS PodsRCS Pods
WingFwd Exposed WingAft Exposed WingElevon
TailFinUpper RudderLower Rudder
Total
Aluminum
21282033871231*8360021*97366032837 21*00
53>*37821*3700
22265UO1*20
1*8,869
Titanium
3611
3611
Non-Metallic
1128U561332665028731*36i*U
6751*75
6869
Total *
(38,lU6)223
110U81*1(61*26?1*U2502781*1*00375387l*21*00
(18,017)601882993700( '3,186)222651*0U20
59,31*9 LBS
* 10$ growth not included
Figure 7 Orblter Structural Arrangement 53
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APPENDIX A
SUMMARY OF INTERSTAGE FORCES
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APPENDIX B
RESONANT FREQUENCY OF BODY
SUSPENDED FROM THE TWO ANGLED WIRES
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OAC 326A REV 2
12-71 1Z5MREPORT AJASl
DATE
35-
CODE zesiz
PAGE S-2
The analysis uses > Liagr ange ' s equations to derive the equationsof motion. The potential energy is calculated from the verticaldisplacement of the center of gravity due to translation and rota-•tion. The kinetic energy is determined by adding t.he. aompojient sdue to linear and rotational velocities. Assumptions of smallmotions and equality of small angles are used to simplify the i- . \.~analysis. The longitudinal flexibility due to the air springsj. £#4the cable elasticity-are also omitted as further simplifying,assumptions. '" " '
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