Hot ruddervator apparatus and method for an aerospacecraft

TECHNICAL FIELD

This invention relates to airfoils for aerospacecraft, and more particularly to a ruddervator for an aerospacecraft which comprises a hybrid construction that enables lower production costs and improved life and reliability of the ruddervator.

BACKGROUND OF THE INVENTION

Current control surfaces for advanced aerospacecraft are formed by a carbon-based ceramic matrix composite (CMC) hot structure with conventional rib-stiffened structure and a mechanically fastened skin. The X-37 aerospacecraft presently in use incorporates a control surface termed a “ruddervator” with the above-described construction which makes use of carbon/silicon carbide (C/SiC). This construction is shown in FIG.

1

. The mechanically fastened upper skin

10

is secured by a high temperature metal, ceramic or ceramic composite fasteners at locations

12

to an integral C/SiC lower skin and substructure

14

. A C/SiC tail tip

16

is used to close the end of the ruddervator. A titanium spindle

17

is used to rotate the ruddervator as needed. Thermal protection system seals

18

,

20

and ring

22

are used to help mount the ruddervator to the fuselage of the aerospacecraft.

The X-37 ruddervator approach described above uses an expensive 2800° F. CMC system in a 2400° F. “hot structure” application and uses an aircraft-like structural approach at the elevated temperature. The term “hot structure” refers to the temperature of the primary load-carrying structure, in this case the CMC and supports used at 2400° F. This construction reduces the service life of the fasteners. Furthermore, carbon-based CMCs generally require complex and costly tooling, unique and expensive infiltration/furnace facilities, and fabrication cycles of six months or more. The use of new materials under development, such as oxide fibers/oxide matrix-based CMC (Oxide-CMC), provide opportunities to design control surfaces in more cost-effective ways including, but not limited to, maintaining internal supports and attachments below 600° F.

For present and planned reusable hypersonic vehicles there are also size constraints on control surfaces due to available volume which restrict the use of conventional, lower cost structure insulated with bonded tile thermal protection. The current solution is to use the CMC for control surface hot structure in areas which do not require their extreme high temperature properties. The result is high initial and recurring costs for these parts as well as weight penalties at high part counts. Without an order of magnitude reduction in thermal structure costs, commercial reusable access to space will be difficult, if not impossible, to achieve.

It is therefore a principal object of the present invention to provide a new construction for a ruddervator for an aerospacecraft which can be produced more inexpensively from a simpler fabrication process, and which has improved life and reliability over the conventional mechanically fastened upper skin-to-substructure construction presently in use for ruddervator applications.

It is another object of the present invention to provide a hybrid control surface for an aerospacecraft which can be manufactured more economically, which is simpler to repair, and which does not make use of typical mechanical fasteners to secure an upper skin to a substructure.

It is still another object of the present invention to provide a ruddervator for an aerospacecraft having a simplified design which requires significantly fewer independent component parts being needed for the construction of the ruddervator.

SUMMARY OF THE INVENTION

The above and other objects are provided by an airfoil for an aerospacecraft. The airfoil comprises a ruddervator having a oxide fiber/oxide matrix-based ceramic matrix composite (oxide-CMC) fabric which is secured to a ceramic foam insulation in the shape of an airfoil section when viewed cord-wise. A plurality of airfoil sections attached adjacent to one another form the ruddervator.

The oxide-CMC fabric is fused over the rigid ceramic foam insulation. The rigid foam insulation includes a hollowed out area through which at least one frame element extends. The hollowed out area of each airfoil section includes a plurality of integrally formed securing members, which in the preferred embodiment comprise lugs, which are secured to structure on the frame element. In one preferred form the frame element comprises a titanium box beam.

Each of the airfoil sections are secured to the frame element such that a lower end of one panel is positioned nestably within an upper end of its lower adjacent airfoil section. The frame element is secured to a torque box of the aerospacecraft such that the entire airfoil can be rotated as needed during flight.

The airfoil of the present invention thus does not require mechanical fasteners to be used to secure a skin to an independent substructure. The construction of the present invention further serves to reduce the cost and weight of the airfoil in large part because of the lower cost, higher specific strength and stiffness of the materials employed. Cost is also reduced because with the common design of the nesting airfoil sections, a single female lay-up mold can be used for the fabrication of all of the oxide-CMC fabric/ceramic foam insulation airfoil sections. The manufacturing cost is further reduced by utilizing the reduced tooling complexities of oxide-CMC fabrication processes over CMC fabrication processes.

BRIEF DESCRIPTION OF THE DRAWINGS

The various advantages of the present invention will become apparent to one skilled in the art by reading the following specification and subjoined claims and by referencing the following drawings in which:

FIG. 1

is a perspective view of a prior art construction of a ruddervator for an aerospacecraft;

FIG. 2

is a side view of a ruddervator for an aerospacecraft in accordance with a preferred embodiment of the present invention;

FIG. 3

is a cross sectional end view of one airfoil section of the ruddervator of

FIG. 2

taken in accordance with section line

3

—

3

in

FIG. 2

;

FIG. 4

is a cross sectional end view of one airfoil section taken in accordance with section line

4

—

4

in

FIG. 2

;

FIG. 5

is a cross sectional side view of one airfoil section taken in accordance with section line

5

—

5

in

FIG. 4

;

FIG. 6

is an enlarged view of the area at which the lugs of the rigid foam insulation are secured to one of the titanium box beams forming the frame system of the ruddervator;

FIG. 7

is a side, partial cross-sectional view of the area shown in

FIG. 6

, taken along section line

7

—

7

in

FIG. 6

; and

FIG. 8

is a side cross-sectional view of just the tip close-out panel secured to its adjacent airfoil section.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

Referring to

FIG. 1

, there is shown a ruddervator

100

for a hypersonic space vehicle such as an aerospacecraft

101

. The ruddervator

100

is used to help steer the aerospacecraft during flight. The ruddervator

100

includes a plurality of identical airfoil sections

102

a

-

102

d

which are secured adjacent to one another in a nesting fashion. Airfoil section

102

a

is disposed nestably within a oxide-CMC laminate closeout panel and thermal barrier retainer

104

which acts as both a lower overlapping close out seal and as part of the retainer for a standard flexible thermal barrier

106

. At the upper end of the ruddervator

100

a specially designed oxide-CMC panel incorporates a close-out foam sandwich cap

108

which is secured to airfoil section

102

d,

and which closes off the upper end of the ruddervator

100

.

Referring to

FIGS. 2-4

, a pair of elongated, highly rigid and temperature resistant frame elements

110

are disposed in parallel relationship to one another and used for supporting each of the airfoil sections

102

a

-

102

d,

the close-out panel

108

, and the lower close-out panel and thermal barrier retainer

104

. With specific reference to

FIG. 3

, it will be noted that each of the airfoil sections

102

a

-

102

d

and the lower close-out panel and thermal barrier retainer

104

each include an oxide-CMC panel

112

which is comprised of preferably multiple plies of oxide-CMC fabric. The oxide-CMC panel

112

is fused over a substrate of rigid, ceramic foam insulation

114

. The topmost outer mold line (OML) ply of the oxide-CMC panel

112

of each airfoil

102

is infused with a high-emissivity (i.e., black) coating, such as reactive cured glass (RCG) or silicon carbide, to provide plasma heating re-radiation outward to reduce internal temperatures of the airfoil section

102

. The ceramic foam insulation

114

and the oxide-CMC panel

112

comprise an airfoil shape when viewed cord-wise. The ceramic foam insulation

114

includes a hollowed out section

114

a

through which the frame elements

110

extend. Integrally formed on an interior surface

116

of the ceramic foam insulation

114

is a plurality of lugs

118

. In the preferred embodiment four such lugs

118

are formed on the interior surface

116

of the ceramic foam insulation

114

. A pair of devises

120

are secured to each frame element

110

to allow each airfoil section

102

to be secured to the frame elements

110

.

Referring further to

FIGS. 2 and 3

, the frame elements

110

preferably comprise titanium box beam members. A lower end

110

a

of each frame element

110

is secured to a torque box

122

(FIG.

2

). This box

122

consists of a honeycomb sandwich-panel construction with access panels

124

included preferably in at least two locations on the torque box

122

. Carrier panel

126

is provided to allow access to the torque box

122

and actuator spindle

125

which controls movement of the ruddervator

100

. Preferably four carrier panels

126

are provided to allow access to the interior of the torque box

122

. The torque box

122

is covered by a plurality of RCG toughened, uni-piece, fibrous insulation (TUFI) coated alumina-enhanced thermal barrier (AETB) tiles

128

. Internal fittings (not shown) within the torque box

122

permit attachment of the frame elements

110

to the torque box

122

as well as to the fuselage actuator spindle

125

. It will be appreciated that the structure comprising components

122

-

128

is known in the art and therefore has not been described in extensive detail.

With brief reference to

FIGS. 6 and 7

, each clevis

120

can be seen to include a pair of L-shaped plates

121

a

and

121

b

disposed in facing relation to one another and secured such as by rivets or any other suitable means to the frame element

110

. One lug

118

is secured to each clevis

120

via a threaded nut and bolt fastener

130

. With specific reference to

FIG. 3

, clevis

120

a

preferably includes a small hole

132

, while devises

120

b

and

120

c

include oversized holes

134

. Clevis

120

d

preferably includes a slotted hole

136

. The oversized holes

134

and the slotted hole

136

allow for thermal expansion of the titanium box beam frame elements

110

. The devises

120

a

-

120

d

may comprise either inconel or titanium depending upon the temperatures expected to be encountered during flight.

Referring to

FIGS. 4 and 5

, the two titanium box beam frame elements

110

can be seen to be secured in parallel relationship by a pair of plates

135

. Preferably, a plurality of pairs of plates

135

are riveted or otherwise secured to the frame elements

110

at spaced apart locations along the frame elements. The plates

135

are preferably placed directly behind the oxide-CMC panel

112

where the re-radiation heating is the lowest to limit the surface temperature of the plates

135

. The airfoil

102

b

can also be seen to include an internal shoulder

137

at an upper end

102

b

1

thereof, and a step portion

139

at a lower end

102

b

2

. The shoulder portion

137

receives the step portion

139

of the adjacent airfoil section (i.e., airfoil section

102

c,

in this instance). Step portion

139

fits within the shoulder portion

137

of its lower adjacent airfoil section

102

(in this instance airfoil section

102

a

). The devises

120

permit the nested airfoil sections

102

to be pre-loaded to reduce inter-panel rubbing at the overlap areas of each panel during vibration and plasma leaks from thermal expansion. The overlap and pre-load also allow the stiffener flange of one panel

112

to help structurally stabilize the unstiffened end of the adjacent airfoil section

102

.

During construction, the titanium box beam frame elements

110

and the plates

135

are assembled together first. The tip close-out panel

108

is the first component to be mounted onto the frame elements

110

. This is illustrated in FIG.

8

. Next the airfoil sections

102

are individually mounted by passing them over the lower end

110

a

of each of the frame elements

110

and positioning them over the previously mounted airfoil section

102

. During the attachment of each airfoil section

102

, the pre-load is applied at the devises

120

. Once each of the airfoil sections

102

and the close-out panel

108

and thermal barrier retainer

104

are disposed over the frame elements

110

, the entire assembly is hoisted onto the torque box

122

and attached via the access panel

124

in the torque box.

The application of oxide-CMC with the ruddervator

100

of the present invention serves to minimize fabrication costs by using a common design for each of the nesting airfoil sections

102

so that a single and common female lay-up mold can be used for fabricating each of the airfoil sections

102

. The ruddervator

100

forms a lower cost and lower weight control element for an aerospacecraft and can be formed through even simpler fabrication processes, as well as maintained through less complex repair processes. Eliminating the need for oxidation protection coatings also serves to improve the life and reliability of the ruddervator

100

. The construction techniques described in connection with the ruddervator

100

could also be used to form various hardware components for an aerospacecraft such as, but not limited to, elevon flipper doors, vent doors, external tank disconnect arrowhead panels and wing leading edge panels in the lower temperature zones of the vehicle.

Those skilled in the art can now appreciate from the foregoing description that the broad teachings of the present invention can be implemented in a variety of forms. Therefore, while this invention has been described in connection with particular examples thereof, the true scope of the invention should not be so limited since other modifications will become apparent to the skilled practitioner upon a study of the drawings, specification and following claims.