The orbiter elevon seal panel mechanism controls the position of fairing panels between the orbiter wing and elevon. Early mechanism designs used linkages which approximately matched the panel motion to elevon position, depending on panel deflections to maintain sealing. These linkages were refined during orbiter development to match panel motion to elevon motion more exactly, thus reducing panel deflections, loads, and weight. Changes to the adjacent cove seal resulted in the use of curved tension compression links. Mechanism temperatures up to 750 F (locally) posed difficulties in bearing lubrication. Despite the adverse effect of the many fabrication tolerances, the system successfully prevented the entry of 1200 F hot gases into the wing/elevon joint

Ripley, J. G.

English

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SPACE SHUTTLE ELEVON SEAL PANEL MECHANISM
John G. Ripley
Rockwell International
Downey, California
ABSTRACT
The orbiter elevon seal panel mechanism controls the position of fairing panels between the
orbiter wing and elevon. Early mechanism designs use_ linkages which approximately matched the panel
mo_ion to elevon position, depending on panel deflections to maintain sealing. These linkages were
refined during orbiter development to match panel motion to elevon motion more exactly, thus reducing
panel deflections, loads, and weight. Changers to the adjacent cove seal resulted in the use of
curved tension-compression links. Mechanism temperatures up to 750F (locally) posed difficulties in
bearing lubrication. Despite the adverse effect of the many fabrication tolerances, the system has
successfully prevented the entry of 1200F hot gases into the wing/elevon joint.
INTRODUCTION
Development of a space vehicle capable of both aerodynamic flight and entry into the atmosphere
posed a number of challenges in sealing the moving control surfaces to prevent the entry of hot gases.
The orbiter elevons, consisting of an inboard and an outboard elevon on each wing, posed a unique
problem ac the upper wing surface. The elevon hinge line was established near the lower wing surface,
so elev0n motion resulted in large changes in the position of the upper elevon forward edge. The
structure and systems installations prevented use of a stationary seal riding on a large skin radius.
The gap between the elevons and wing required sealing sufficient to prevent interior heating by
the hot entry gases, which reach up to 1200F in that area. These high temperatures would damage the
aluminum structure (restricted to 350F maximum), and the hydraulic systems operating the surfaces could
not be allowed to exceed 275F. In addition to thermal sealing, the upper wing/elevon junction
required a fairing to provide aerodynamic smoothness in an area subjec_ to high vibration, pressure
differentials, and buffet/flutter loads.
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The approach to this challenge was initially developed by Gru_man, the major subcontractor for the
orbiter wings. A system of overlapping panels made of high-temperature alloys bridged Ehe gap between
the wing and the elevon. Each panel had rollers on its trailing edge, and the panels were preloaded
to maintain roller contact. Seals between the rollers and at the panel hinges prevented hot gas
ingestion.
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https://ntrs.nasa.gov/search.jsp?R=19850008652 2020-03-22T18:09:26+00:00Z
The use of mechanisms came about in order to maintain contact of the rollers on the elevon
surface. Predicted burst/crush loads, dynamic requirements, and magnitude of movement precluded
spring loading the panels down. Therefore, at each panel, 2 four-bar linkages were used to cause
the motion of the panel to approximately follow that of the elevon.
This initial design resulted in e set of interf&ce points that became fixed for the reminder
of the program. As shown in Figure I, there were 15 panels per wing. A linkage (Figure 2) connected
the elevon to a hellcrank (mou._ted on the rear wing spar) at 17 locations.
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The bellcrank, in turn, :'::! "5:
connected with either one or two panels, depending on the location, for a total of 29 connections to /..:
panels from the 17 bellcranks per wing. In general, the pivot axis of each bellcrank was skewed : .,.,Ci
relative to the elevon and seal panel hinge lines in an attempt to cOmpromise between the latter two, !_'-_ii'"!f.2
which were not parallel, Since the panel motion did not precisely follo_ the elevon motion, the _V'_'V'"-"'
clevis attaching the upper link to the panel was adjustable, and a high preload was put into each _ ;,_ .....
linkage by bowing the panel. This allowed the rollers to remain in contact without perfectly matched ;_ "_ ""
panel-to-elevon motion.
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PROBLEMS WITH INITIAL DESIGN !_.'_.,_
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The initial design was used on OV-IOI (The Enterprise) for the Approach and Landing Tests (ALT). ,,."_---_"
The seals and materials were not fully developed for high-temperature use at that time. The ALT _ 1:71_'i,"
sometimes fall to roll, giving high friction and a jolt at each new motion. The holes in the roller ._.._
clevises elongated under the high loads, changing the adjustment and spoiling the pivot surfaces.
As a temporary measure to allow cOmpletion of the ALT program, the preload on the linkage was
reduced. The bushings in the rollers were replaced by needle-type bearings to reduce friction, _._._r: .
These measures provided acceptable performance in the low temperature ALT program, but the resulting 'i.;if:'_ "'
gaps at the panel trailing edge were not acceptable for operational use. In addition, no anti- :/; ,_-
friction bearings were available that could meet the environmental requirements of the trailing edge _. :'i;.'=._':
rollers. At this point in the development, the mechanism (linkage) weight was approximately 221 ":- 7
pounds per vehicle. _<'_i':.
CHANGES FOR ORBITAL FLIGHT TEST
Following the ALT program, Rockwell International undertook the further development of the
subsystem to meet full orbital mission requirements. Primary goals were weight reduction,
simplification, and a trailing edge configuration compatible with the high entry temperatures. The
approach was to refine the kinematics so that an approximately constant gap would be maintained
between the panel trailing edge and the elevo_ surface throughout the travel. This was envisioned
as eliminating the rollers and reducing loads (a high preload would no longer be required to maintain
panel trailing edge contact). A new metal seal configuration was developed by the thermal protection
system designers to have greater initial deflection and thus tolerate greater variations in trailing
edge gap. The lower loads, combined with use of lightweight swaged-end aluminum rod assemblies, gave
an anticipated weight savings of 114 pounds per vehicle (52%).
TECHNIOUES USED
The challenge now faced by the designers was the sheer magnitude of the work required.
Seventeen (17) lower stages and 29 upper stages had to be kinematically optimized to maintain the
trailing edge gap within + o060 inch over the elevon travel. Since pivot axes, panel sizes, and
wing contours were unique--at each location, separate kinematic studies were required at each. At
that time (1976), no Computer Aid Design (interactive graphics) facilities were available at
Rockwell.
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The kinematic op_imizatlon became a joint effort of the Mechanical Design Group, with
Mr. I. Awtamonow as lead designer, and the Numerical Design Group, represented by R. Chu, Supervisor,
A. Shapiro, and B. givera. The Numerical Design group's primary responsibility was _he mathematical
modeling of vehicle lines and the prograaning of large computer controlled plotters to generate lines
drawings. Using Chls background, the analysts developed a geometry analysis program for the main
frame digital computer. The program allowed the entry of daEa on the basic pivot points and lines at
each location. The radii and angles of the bellcrank were the only parameters which could be varied
without changing the interface points built into the s_ructure. With the chosen values entered, _he
user could input the elevon angle, and _he program would calculate the positions of all _he joints
and _he size of the _railing edge gap.
This program was used aC each location Co develop the linkage geometry. The designers proceeded
_uch as _hey would have "on the board," using the original ALT linkage as a start and converging on a
satisfactory solution by observing the sensitivity of the gap to each variable ac the elevon up,
neutral, and down positions. By this process of "enlightened trial and error_" _he geometry of all
_he linkages was established in abou_ one year.
DETAIL DESIGN
As a geometry was established a_ each •location, spot layouts were made co assure adequate
clearances. However, _he resources were not available to allow a detailed check layout of each
location. Design of link and bellcrank de_ails followed the establishing of the geometry.
Convential aerospace ligh_ weight design pracEices were followed. The most significant obstacle to
detail design was _he high temperature anticipated at the rod end attaching _o the seal panel. A
rod end using carbon-graphite-MoS2 bearing inserts was specified in order to withstand the
approximately 8O0F temperature predicted. Bearings of the same material, but in a different
configuration, were used in o_her orbiter systems - e.g., the Payload Bay Door Hinges.
The final loads i_era_ion for _he subsystem was completed in 1977, when detail design was well
underway. Analysis revealed rhac _he light links had insufficient stiffness based on flutter/buffet
criteria. As a res_l_, mos_ _ube/link derails were redesigned. The larger diameter and wall thickness
resulEed in a weight increase, but _he subsystem weight remained close Co the Carger weight of 107
pounds.
LATER CHANGES
The elevon cove seal on the lower wing/elevon surface was redesigned in 1977. The diameter of
the seal increased to the poin_ _haC _he lower links would interfere with _he seal as shown in
Figure 3. This problem was discovered in spoC layouts _rior co completion of detail release. Changes
in _he critical seal were nor feasible, so _he lower links were redesigned _o clear the seal. The
resulting curved links added 21 pounds per ve_icle to the subsystem.
The body of the curved links was hollow aluminum, the curved and straight portions being one
conclnuous par_ (Figure 4). The initial fabrication sequence consisted of forming an oversized
machined _ube to the curvature, chemically milling and machining final dimensions, and heat treatment.
Difficulties were encountered due _o the higher rate of chemical material removal in the stretched
(formed) areas as well as dimensional changes induced by heaC _rea_men_. Manufacturing technique
developmen_ eventually produced trouble free fabrication. The sequence evolved into the following:
machining _he scraigh_ _ube Co final dimensions; forming the _ube _o a smaller radius of curvature
than _he final requirement; and hea_ _reacmen_, during which _he tube s_raightened ou_ to final
dimensions. The amoun_ of overbending required co achieve the correct curvature after heat treatment
was de_ermined by experiment.
863
At some locations, an additional clearance problem precluded the use of the large diameter
curved tubes, in these areas, the curved portion of the tube entered a channel fitting with the
elevon up (Figure 5). An alternate design was developed as shown in Figure 6. The narrower section
and anti-rotation lugs provided the additional clearance required. However, the greater weight
of these links resulted in the total subsystem weight increasing to 196 pounds per vehicle.
These heavier Inconel 718 links were used at seven of the nine inboard lower stages.
BEARING MATERIAL CHANGE
As described above, the high temperature end of the upper links used a carbon-graphite-MoS2
bearing insert material. This material will withstand temperatures up to 2500F. However, the
manufacturer of the bearings found that the spherical ball rod end with carbon-graphite bore
presented unanticipated fabrication problems. Cracking of the insert material during fabrication
resulted in an extremely high rejection rate. Virtually no bearings were delivered 1-1/2 years afte#
the first order.
In mid-1978, the decision was made to give up on t_e carbon-graphite inserts. A rod end using
a ceramic MoS2 dry film lubricant was specified instead. The new lubricant was restricted to s
maximum of 750F for short duration only, which corresponded to the expected environment with no
margin. The dry film lubricant was therefore being pushed to perform at its maximum capability.
However, experience had been gained by this time with this dry film lubricant in other orbiter
mechanisms, resulting in confidence that the life would be adequate.
INITIAL INSTALLATION PROBLEMS
The installation and adjustment of the first operational system occurred on OV-102 (Columbia).
Owing to the simplicity of the hardware and its similarity to other orbiter mechanisms, a full
qualification program was not performed. Consequently, a number of problems were encountered on
the first article. The primary problem was an inability to hold the trailing edge gap tolerances
originally targeted. The buildup of fabrication tolerances was enough to force the gap tolerance
to increase from ÷ 0.060 (original) to I 0.130 (neutral position), ÷ 0.160 (elevon down), and
+ .215 (elevon upS. The tolerances involved included those on the _levon moldline (contour), panel
_etails, bellcranks, wing spar fittings, elevon hingeline, and elevon clevis fittings. Adjustments
of the link lengths (upper and lower) from nominal were made based on calculated sensitivity until
the best results were achieved. Fortunately, the seal design had enough margin to absorb the
increased gap variations.
Other first article problems included low clearances to hydraulic lines and fasteners, and rubbing
of rod end locking devices against adjacent rod bodies. The problems were considered minor for the
first installation of a system without complete check layouts or mockups. R. Holt of
Rockwell International was the responsible enginecer during the subsystem installation phase.
Installation of 0V-099 (Challenger) encountered very few problems. However, many trailing edge
gaps were at the low side of the tolerance band, and excessive seal wear was noted on many seals
following elevon tests. The damage was found to be caused by rubbing of an area of the seal not
intended to contact. A change in the seal mounting geometry solved this problem and avoided any
change to the mechanism or the seal itself.
SUBSYSTEM PER.FORY.ANCE SUMMARY
Figure 7 illustrates a typical linkage of the final configuration. The mechanism performance
hag been satisfactory on all orbital flights. No functional or leakage problems have been
encountered. The final design, as developed from the ALT version, meets the original challenge
in the intended way - simply and directly.
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Space shuttle elevon seal panel mechanism

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