A hexagonal protective and high temperature resistant system for the Space Shuttle Orbiter consists of a multiplicity of pockets formed by hexagonally oriented spacer bars secured on the vehicle substructure. A packing of low density insulating batt material 18 in each pocket, and a thin protective panel of laterally resilient advanced carbon-carbon material surmounting the peripherals bars and packing. Each panel has three stepped or offset lips on contiguous edges. At the center of each pocket is a fully insulated stanchion secured to and connecting the substructure and panel for flexing the panel toward the substructure and thereby prestressing the panel and forcing the panel edges firmly against the spacer bars

Dunn, T. J.

English

NASA Technical Reports Server

nited States Patent [191 
Dunn 
11 11 4,439,968 
1451 Apr. 3, 1984 
[54] PRE-STRESSED THERMAL PROTECTION 
SYSTEMS 
[75] Inventor: Thomas J. Dunn, Pasadena, Tex. 
[73] Assignee: The United States of America as 
represented by the Administrator of 
the National Aeronautics and Space 
Administration, Washington, D.C. 
[21] Appl. NO.: 418,137 
[22] Filed: Sep. 15, 1982 
[51] Int. ( 3 . 3  ........................... E04B 1/74; B32B 9/04 
[58] Field of Search ......................... 52/404, 506, 541; 
[52] U.S. Cl. ........................................ 52/404; 52/506; 
244/158 A 
244/158 A, 160 
~561 References Cited 
U.S. PATENT DOCUMENTS 
5,091 5/1847 Stewart ............................. 52/54] X 
4,037,751 7/1977 Miller et al. ...................... 52/404 X 
4,220,685 9/1980 Markow et al. .................. 52/404 X 
4,290,250 9/1981 Kusenda ................................ 52/404 
320,822 6/1885 Simons .................................. 52/54] 
4,124,732 11/1978 Leger .............................. 244/158 A 
4,151,800 VI979 Dotts .............................. 244/158 A 
4,280,484 7/1981 Mancosy ........................... 52/506 X 
4,379,381 4/1983 Holcombe ............................. 52/404 
4,379,382 4/1983 Sauder .................................. 52/506 
Primary Examiner-J. Karl Bell 
Attorney, Agent, or Firm-Russell E. Schlorff; John R. 
Manning; Marvin F. Matthews 
m1 ABSTRACX 
A hexagonal protective and high temperature resistant 
system for the Space Shuttle Orbiter consists of a multi- 
plicity of pockets formed by hexagonally oriented 
spacer bars 11 secured ot the vehicle substructure 12, a 
packing of low density insulating batt material 18 in 
each pocket, and a thin protective panel 19 of laterally 
resilient advanced carbon-carbon material surmounting 
the peripheral bars 11 and packing 18. Each panel 19 has 
three stepped or offset lips on contiguous edges, as M-1, 
M-2, M-3, the other three edges, as M-4, M-5, M-6, 
being non-stepped, Le., flat so that complementary 
edges of abutting panels rest against the peripheral bars. 
At the center of each pocket is a fully insulated stan- 
chion 15 secured to and connecting the substructure 
and panel for flexing the panel toward the substructure 
and thereby prestressing the panel and forcing the panel 
edges firmly against the spacer bars. 
9 Claims, 4 Drawing Figures 
1 3  
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FIG.3 
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damage. Also, the silane bums out, therefore the tile 
system must be refurbished between flights to repair 
damage and provide new water repellency. 
An effort has been made to find other materials which 
5 have stronger surface properties, and require a mini- 
mum of refurbishment between flights. One approach is 
to use high temperature resistant resilient material, such 
as advanced carbon-carbon, as outer panels for the large 
high temperature area on the underside of the wings and 
Space Shuttle Orbiter experiences stresses from a wide 
variety of loading conditions, such as mechanical, aero- 
dynamic, and vibro-acoustic, as well as extreme temper- 
The present invention relates to a thermal protection ature variations, it is necessary that the panels used for 
system, and more particularly to a reusable thermal l5 thermal protection be of sufficient strength and at- 
protection system for a portion of the Space Shuttle tached to the substructure in such a manner as to be able 
Orbiter. to withstand the stresses and loads of space flight and 
Space vehicles are subjected to increases in tempera- that, at the Same time, the system provide the required 
ture during ascent and reentry. A heat shield is used to thermal properties without a weight penalty. 
protect the metallic structure of the vehicle during such 2o Consequently, an object of the present invention is to 
periods. During the early space programs, manned ve- develop a thermal protection system of protective pan- 
hicles were used for single trips and the heat shields els for the surfaces of the Space Shuttle Orbiter subject 
were formed of ablative material. However, with the to high temperature, which will withstand the stresses 
advent of the Space Transportation System which used and loads of repetitive space flights with a minimum of 
a Space Shuttle Orbiter designed to make numerous 25 refurbishment. 
trips into space, it was necessary to develop a reusable Another object is to provide improved attaching 
thermal protection system. The Space Shuttle Orbiter is means for a thermal protection system of protective 
in some respects Similar to a large airplane and conse- panels in which the attachment means are insulated to 
quently various areas of the vehicle are subject to vary- minimize oxidation from heating. 
ing surface temperature extremes. For example, the 30 A further object is to provide a thermal protection 
nose cone and leading edges of the wings are subject to system of hexagonal protective panels. 
the greatest heating and require a system which will Still another object is to improve the impact resis- 
F. (181 1" K.). A coated reinforced carbon-carbon com- 
posite material (RCC) has been found satisfactory for 35 
these areas. On the Other hand* the cargo bay doors are The reusable thermal protection system of the present 
which will provide protection for surface temperatures carbon polygonal protective panels attached to the 
ble reusable surface insulation (FRSI) fully described in a chanically attached insulated stanchion and a plurality 
which will provide protection for surface temperatures 
only require protection for surface temperatures up to 
1200" F. (9220 K.). The high temperature system is 
PRE-STRESSED THERMAL PROTECTION 
SYSTEMS 
DESCRIPTION 
Origin of the Invention-Field of the Invention 
The invention described herein was made by an em- 
Of the United States Government and may be 
manufactured and used for the government of the 
payment of any royalties thereon or therefor. 
United States for governmental purposes without the lo Since the thermal protection system Of the 
BACKGROUND OF THE INVENTION 
provide protection for surface temperatures UP to 2800" tance and durability of the thermal protection system. 
SUMMARY OF THE INVENTION 
subject to the least heating and Only require a system invention is formed of a plurality of advanced carbon- 
aluminum substructure of the vehicle through a me- to 700" F* (644" K*)a It has been found that the 
U.S. Pat. No. 4,151,800 is satisfactory for this area. The of adhesively attached insulating spacer bars. Contigu- undersides Of the wings and require a system ous pockets co-extensive with the surface to be pro- 
tected are formed. Batt insulation is packed in each 
bars and a hexagonal carbon-carbon panel bears against 
to 2300" F* (15330 K*) the sides Of the 
45 pocket beween the stanchion and the peripheral spacer 
known as high temperature resuable surface insulation each batt unit with its periphery resting against 
(HRSI) and the low temperature system as low temper- the framework Of surrounding spacer bars* The panels 
ature reusable surface insulation (LRSI). Presently, are initially, conically dished (convex each 
ceramic tiles cut from rigidized blocks of highly refined, 50 having its Periphery foIl'ned by three contiguous 
chemically pure, silica fibers are used for such systems. stepped lips, the intervening three edges being un- 
The material, known as LI-900, is manufactured by 
Lockheed ~ i ~ ~ l ~  and space Corporation. The tiles are 
attached to the orbiter by the attachment arrangement 
fully described in U.S. Pat. No. 4,124,733. 
For areas around openings and other locations requir- 
ing greater mechanical strength, a denser, stronger sil- 
ica material described in US. Pat. No. 3.952.082. and 
stepped, i.e. flat. Insulated fastener means flex the panel 
laterally into substantially flat configuration, thereby 
Prestressing the Panel and pressing the Peripheral edges 
55 thereof into firm contact with the spacer bars against 
which they rest. Due to the hexagonal Pattern of the 
COver Panels, no long, Unbroken straight lines Occur 
between the abutting edges of adjacent Panels. 
, ,  
known as LI-2200, is used. In addition, a new fibrous 
refactorv comDosite insulation material. described in 60 BRIEF DESCRIPTION OF THE DRAWINGS 
U.S. Pat. No. i,148,962, and known as FRCI, is being 
considered for use. The tiles of all three materials are 
coated with a reaction cured borosilicate glass coating 
made in accordance with U.S. Pat. No. 4,093,771. Fur- 
ther, the tiles are provided with a deposition of a silane 65 thermal protection system of the present invention. 
to provide water repellency. However, the coating as 
well as the tile material itself is fragile and has a low 
impact resistance, and, therefore, is easily subject to 
In the accompanying drawings: 
FIG. 1 is a perspective view of the Space Shuttle 
Orbiter viewed from the underside. 
FIG. 2 is a transverse cross section of a portion of the 
FIG. 3 is a plan view of a portion of the thermal 
protection system, parts being broken away to show 
underlying structure. 
3 
4,439,968 
FIG. 4 is a detail section taken on broken line 4-4, of 
FIG. 3. 
DETAILED DESCRIPTION O F  THE 
PREFERRED EMBODIMENT 
The Space Shuttle Orbiter, as shown in FIG. 1, is 
provided with various reusable thermal protection sys- 
tems, according to heat resisting requirements. The 
nose cone 2 and leading edges 3 of the wings 4, are 
subject to the greatest amounts of heat on reentry and 
are protected by reinforced carbon-carbon. The bot- 
toms of the fuselage and wings are subject to tempera- 
tures up to 2300" F. and are provided with the high 
temperature thermal protection of the present inven- 
tion, generally designated 6. The sides 7 of the fuselage 
and the tail 8 are provided with low temperature reus- 
able surface insulation. The cargo bay doors 9, subject 
only to temperatures up to 700" F., are provided with 
flexible, reusable surface insulation. 
FIG. 2 illustrates in detail one of the numerous high 
temperature resistant units making up the protection 
system 6. At the edges of the unit are spacer bars 11 
formed of FRCI-12 which are adhesively secured to the 
vehicle substructure 12 through thin strips of Nomex 
felt 13 by thin layers of RTV-560 adhesive 14. The 
spacer bars are disposed to form hexagonal pockets in 
the center of each of which there is an insulated stan- 
chion 15 having a cylindrical base 16 of FRCI-12 me- 
chanically attached to substructure 12. Extending 
above base 16 of the stanchion is a spool 24 of a high 
temperature resistant, super alloy, such as alloy 188. 
The spool stem 25 is surrounded by an annular plug 27 
of high temperature resistant, ceramic insulation such as 
LI-900. Fasteners 17 extend through the base plate of 
spool 24, as well as base 16, to mechanically secure 
stanchion 15 to substructure 12. The pockets, between 
the side bars and stanchion, are packed with batt insula- 
tion 18 of low density, high heat resistant material, such 
as alumina-silica fibers, in nickel foil. Each batt is sur- 
mounted by a hexagonal protective panel 19 of ad- 
vanced carbon-carbon having a depressed central hub 
portion 20. 
Each protective panel 19 is formed of high tempera- 
ture resistant, laterally resilient coated, twelve-ply, ad- 
vanced carbon-carbon, the assembly having a final 
thickness of approximately 0.140 inch. The panel is 
initially formed in an upwardly convexed shallow coni- 
cal or dished shape with hub 20 at its apex (dot dash 
lines 28). A fastener 29, extending through aligned bores 
in hub 20 and spool flange 26, has its shank threadedly 
engaging internal threads in spool stem 25. A broad 
head 30 engages hub 20 so that rotation of fastener 29 
will draw hub 20 against stanchion 15 to flatten and 
prestress panel 19 and draw its peripheral edges tightly 
against spacer bars 11. In the installed position, the 
recess formed by hub 20 is closed by a disc 32 of ad- 
vanced carbon-carbon, sealed in position flush with the 
surrounding panel by ceramic cement injected in an 
annular groove 33 about the disc. Such construction 
provides fastener 29 with some protection from the heat 
of reentry and slows down oxidation. 
FIGS. 2, 3, and 4 illustrate the complementary mat- 
ing edges of one protective unit M and six surrounding, 
contiguous, indentical units N, 0, P, Q, R and S ,  
whereby tight joints are formed between the hexagonal 
panels and spacer bars. Each panel has three stepped or 
offset edge lips, as M-1, M-2, and M-3, the other three 
edges, M-4, M-5, and M-6, being unstepped or flat and 
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complementary to the offset lips. In the assembly, each 
stepped lip underlies an unstepped or flat abutting edge 
of an adjacent panel and the complementary overlap- 
ping lips rest firmly against the corresponding spacer 
bar. FIG. 4 shows the interleaved bar edges at one 
corner whereof. FIG. 3 shows portions of the protec- 
tive panels 0 and P cut away to illustrate the relation- 
ship of the underlying lips. It is noted that the exposed 
corner of lip P-1 fits between lips M-2 and M-3 which 
are joined by the integral, offset gusset 34 below lip P-1 
for support and aligment purposes. While the intersec- 
tion between lips M-2 and M-3 and gusset 34 are ogee 
curves, as shown in FIG. 4, these curves appear as 
dotted lines in plan view FIG. 3, for simplicity. The 
arrangement of FIG. 4 is used wherever three panel lips 
come together in the system. 
The installed panels are prestressed in bending, the 
fasteners in tension and the spacer bars in compression 
so that the edges of the panels remain in firm contact 
with overlapped edges of the adjacent panels. More- 
over, the thickness, size and precurvature are sized so 
that the forces generated on lift-off will not overcome 
the preload. 
It has been found that having the panels hexagonal in 
planform permits identically shaped panels to mesh 
together in a smooth array where each panel has the 
greatest possible ratio of area to circumference. 
Accordingly, there is provided a thermal protection 
system for the Space Shuttle Orbiter which is weight 
competitive with the HRSI material presently used for 
the same purpose. The advanced carbon-carbon panels 
will provide a durable and impact resistant surface and 
smoother outer mold line with good maintainability and 
minimum turn around time and effort. 
The exclusive use of all modifications as come within 
the scope of the appended claims is contemplated. 
I claim: 
1. A reusable high temperature, thermal protection 
system for a vehicle substructure comprising a plurality 
of spacer bars secured in pocket forming groups on the 
vehicle substructure, packing of low density insulation 
in each pocket, an outwardly dished protective panel of 
high temperature resistant, resilient material surmount- 
ing each pocket, and means flexing the central portion 
of each panel toward said substructure for prestressing 
each panel and pressing the periphery thereof firmly 
against the surrounding spacer bars. 
2. A thermal protection system as described in claim 
1 in which said flexing means comprises a stanchion 
mechanically connected to the substructure and in- 
cludes threaded means for laterally flexing said panel. 
3. A thermal protection system as described in claim 
1 including a plurality of said panels forming a continu- 
ous surface and in which each of said panels is hexago- 
nal whereby abuttments of said panels form nonstraight 
lines across the system. 
4. A thermal protection system as described in claim 
2 in which each of said panels is made of advanced 
carbon-carbon material. 
5. A thermal protection system as described in claim 
2 in which each of said panels is hexagonal and has three 
contiguous edges forming offset lips and the other three 
edges are flat forming overlapping abutting structure. 
6. A thermal protection system as described in claim 
4 including insulation material fully surrounding said 
stanchion. 
7. A reusable thermal protection system for a sub- 
structure of a space vehicle comprising a plurality of 
5 
4,439,968 
6 
groups of hexagonally arranged and pocket forming 
spacer bars secured to said substructure, laterally resil- 
ient, conical outer protective panels with edges resting 
against each of said groups of bars, three contiguous 
ones of said edge portions forming offset lips and the 
other three being flat for interfitting with complemen- 
tary edges of abutting panels, a packing of low density 
matt insulation in each of said pockets and an insulated 8 
stanchion intermediately connecting each Panel and 10 in which the head of the fastener is covered with high 
said substructure and having means for centrally flexing 
panel and firmly press the edges thereto against said 
bars. 
set forth in claim 7 
wherein the stanchion is mechanically attached to the 
5 substructure and the means flexing said panels is formed 
of a threaded fastener extending through the central 
portion of each panel and engaged with threaded 
in the stanchion. 
9. The thermal protection system set forth in 
8* The thermal Protection 
temperature resistant material. 
each panel toward said substructure to prestress each * * * * *  
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Pre-stressed thermal protection systems

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