Survivable and reusable launch vehicle

FIELD OF THE INVENTION

The present invention relates in general to air vehicles and more specifically to a reusable weapon delivery air vehicle having an arch wing.

BACKGROUND OF THE INVENTION

Reusable launch vehicles including the space shuttle are known. Weapon delivery systems which are self guided following launch of the system are also known. An example of this type of system includes the cruise missile. The cruise missile is normally rocket launched from a stationary or mobile platform and includes its own internal navigation equipment to enable the cruise missile, given its originating location coordinates, to identify and fly under its own power to a specific target. The cruise missile is very effective at delivering relatively small explosive payloads to a target. Disadvantages of the cruise missile include: (1) it travels at subsonic velocities and is susceptible to being detected and destroyed by enemy fire; (2) it includes its own engine and fuel, reducing its payload; (3) it can strike against a single target only; and (4) it is not a reusable platform.

It is therefore desirable to provide a reusable air vehicle having improved survivability and a reduced detection signature, which carries no propulsion system, but is operable at high altitudes and travels at higher than sonic speed. It is also desirable to provide a reusable air vehicle to reduce the operating costs of delivering a weapons payload.

SUMMARY OF THE INVENTION

According to a preferred embodiment of the present invention, a Mach-velocity reusable launch vehicle (RLV) is provided which has an independently propelled launch phase, a powerless glide phase, a weapons release phase, and a landing phase. The RLV includes a generally tubular shaped body having an aft end and a forward end, and a payload section. An arch wing is supported by the body aft end. The arch wing has an upper and a lower wing joined at distal ends by two curved end plates. A forward steering device is provided as a pair of flight control surfaces mounted on opposed sides of the RLV at about the horizontal centerline. A nose assembly is connected at the RLV forward end having an upward-directed, fixed angle-of-attack to generate forward end lift. Radar absorptive and translucent materials are used throughout the RLV. A rear-launch system is preferably provided to discharge a weapons payload.

The RLV of the present invention is preferably launched as known in the art by attachment to a rocket propelled vehicle capable of releasing the RLV at hypersonic velocity (i.e., Mach 3 and higher) and to an altitude of 100,000 ft or greater. The RLV achieves maximum elevation and velocity during the launch phase, disengages from the launch platform, reenters the atmosphere (if necessary) at a predermined angle-of-attack, and thereafter travels to a pre-designated location in a powerless glide phase. In a preferred embodiment, to reduce the chance of detection and to permit weapons release at more than one location, the RLV maintains hypersonic velocity during the glide phase and the weapons release phase by discharging its payload from the aft end. Following the weapons release phase, and during the landing phase the RLV is steered toward a landing area remote from the weapons release location(s). Landing gear extend from the underwing for the landing phase.

Thermal insulation attached under the body and the lower wing under-side absorbs/dissipates heat generated during RLV atmospheric re-entry. A high angle-of-attack is maintained and the temperature generated by the hypersonic velocity of air moving under the RLV is dissipated by the thermal insulation materials which protects the vehicle's lower surfaces and shields the upper body and the upper wing of the arch wing from the airstream's thermal impact.

Similar features to the arch wing, the forward steering device, and the landing gear are also described in U.S. patent application Ser. No. 10/200,692, filed Jul. 22, 2002, which is incorporated herein by reference.

Further areas of applicability of the present invention will become apparent from the detailed description provided hereinafter. It should be understood that the detailed description and specific examples are intended for purposes of illustration only and are not intended to limit the scope of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention will become more fully understood from the detailed description and the accompanying drawings, wherein:

FIG. 1

is a perspective view of a preferred embodiment of a reusable launch vehicle of the present invention;

FIG. 2

is a plan view of the reusable launch vehicle shown in

FIG. 1

;

FIG. 3

is an end elevation view taken along directional line

3

—

3

of

FIG. 2

identifying the general shape of the arch wing of the present invention;

FIG. 4

is a side elevation view taken along directional line

4

—

4

of

FIG. 2

showing the nose assembly angle-of-attack angle as well as locations for the payload section and payload release section;

FIG. 5

is a side elevation view similar to

FIG. 4

showing the reusable launch vehicle during a re-entry phase following the vertical launch;

FIG. 6

is a perspective view of another preferred embodiment of the present invention which includes a ring canard and a plurality of ogive nose sections selectable for use on the reusable launch vehicle;

FIG. 7

is a side elevation view similar to

FIG. 5

showing the alternate embodiment ring canards of FIG.

6

and their capability to control pitch of the reusable launch vehicle;

FIG. 8

is a perspective view showing a preferred rear release of payload from the reusable launch vehicle aft end;

FIG. 9

is a perspective view of a preferred embodiment of a steering device of the present invention identifying a standoff clearance permitting boundary layer flow to propagate between the steering device and the external surface of the reusable launch vehicle;

FIG. 10

is a section view taken at Section

10

—

10

of

FIG. 2

, showing radar absorbing/translucent materials applied to the RLV of the present invention;

FIG. 11

is a perspective view of yet another preferred embodiment of the present invention having an inflatable tail cone disposed at an aft end of the reusable launch vehicle; and

FIG. 12

is a section view taken at Section

12

—

12

of

FIG. 3

, showing exemplary flight control surfaces on the trailing edge portions of the lower wing.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The following description of the preferred embodiment(s) is merely exemplary in nature and is in no way intended to limit the invention, its application, or uses.

Referring to

FIG. 1

, in accordance with a preferred embodiment of the present invention, a reusable launch vehicle (RLV)

10

includes a body

12

, an arch wing

14

, and a nose assembly

16

. The body

12

has a generally tubular shape which supports the arch wing

14

at an aft end of the body

12

. The body

12

also supports the nose assembly

16

at a forward end of the body

12

. A forward steering device

18

is mounted on the body

12

and is rotatably disposed such that the steering device

18

creates side thrust at the forward end of the body

12

when displaced from its stowed position shown.

The arch wing

14

includes an upper wing

20

, a lower wing

22

, and a pair of curved end plates

24

joining distal ends of each of the upper wing

20

to the lower wing

22

. Each of the upper wing

20

, the lower wing

22

, and the curved end plates

24

include rounded leading edges

26

and squared trailing edges

28

, respectively.

The upper wing

20

has a swept wing design which will be discussed further in reference to FIG.

2

. An upper wing point-of-contact

30

is formed between the upper wing

20

and the body

12

. From the upper wing point-of-contact

30

, the upper wing leading edges

26

each taper back to a forward facing point of the curved end plates

24

. The upper wing

20

is supported from an upper surface of the body

12

and has a generally planar shape. The lower wing

22

also has a swept wing design which will be discussed further in reference to FIG.

2

. The lower wing

22

has a lower wing point-of-contact

32

at a forward leading edge of the lower wing

22

at a junction with the body

12

. From the lower wing point-of-contact

32

, the lower wing leading edges

26

each taper back to a forward facing point of the curved end plates

24

. Similar to the upper wing

20

, the lower wing

22

is supported from a lower surface of the body

12

.

Referring now to

FIG. 2

, a body aft end

34

extends aft of the upper wing

20

and provides a payload discharge face

36

. The payload carried by the RLV

10

exits in a payload discharge direction A as shown.

As noted above, the upper wing

20

is connected to the body

12

at the upper wing point of contact

30

. The swept wing shape of the upper wing

20

includes a swept wing angle &thgr; measured from a body longitudinal centerline

38

. The swept wing angle &thgr; is fixed for each arch wing design, and can vary between approximately 20° to approximately 60°. The lower wing point of contact

32

is formed at the body

12

along the body longitudinal centerline

38

. The lower wing

22

displaces forward of the upper wing

20

by a lower wing extension length B as shown. The advantage of providing the lower wing extension length B will be discussed in reference to FIG.

5

. The lower wing

22

is also a swept wing having an angle within the range of the swept wing angle &thgr;.

At a forward end of the body

12

the steering devices

18

are mounted on opposed sides of the body

12

. Each of the steering devices

18

rotate at a rotation point

40

to permit a steering device deployment angle &agr; as shown. The steering device deployment angle &agr; forms an angle measured from the body longitudinal centerline and ranges from approximately 0° to approximately 30° when the steering device

18

reaches the steering device deployed position C (shown in phantom).

Referring to

FIG. 3

, the design of the arch wing

14

is provided in greater detail. The upper wing

20

has a generally planar shape horizontally mounted from an upper surface of the body

12

. The curved end plate

24

is provided to join distal ends of both of the upper wing

20

and the lower wing

22

. The curved end plate

24

serves several purposes, including: (1) the curved shape is a structurally efficient form for joining the upper wing

20

to the lower wing

22

; (2) the curved end plate

24

prevents high pressure air from the underside of the lower wing

22

from spilling about the distal end of the lower wing

22

to the upper side of the lower wing

22

; and (3) providing curved surfaces to join the upper wing

20

to the lower wing

22

increases the directional stability of the air vehicle. The lower wing

22

is provided with a horizontal portion adjacent to the lower wing attachment point to the body

12

. Outboard of the horizontal portion of the lower wing

22

, the lower wing

22

includes a dihedral angle &phgr; commonly employed for aircraft wings. The dihedral angle &phgr; ranges from approximately 0° to approximately 15°. In another preferred embodiment (not shown), the upper wing

20

can also include the dihedral angle &phgr;.

FIG. 3

also shows each of the steering devices

18

in their fully deployed position. When deployed, an air flow clearance

42

is provided between the outer surface of the body

12

and an inner surface of each of the steering devices

18

. The air flow clearance

42

provides for high velocity air flow in this standoff region which allows boundary layer air to propagate through the entire length of the steering devices

18

, increasing its steering effectiveness when these devices are deployed.

FIG. 3

also shows a landing gear set

44

in its deployed position. The landing gear set

44

is normally stowed during the launch and the powerless flight phases of the operation of the RLV

10

. The landing gear set

44

can include skis, struts, or wheels. An exemplary quantity of three components are shown for the landing gear set

44

. Two of the components are disposed in the lower wing

22

and a third component is disposed at a lower portion of the forward end of the body

12

. Each component of the landing gear set

44

is provided with a mechanism (not shown) to deploy and retract the landing gear set

44

.

Referring to

FIG. 4

, both the steering device

18

and the nose assembly

16

are shown in greater detail. The nose assembly

16

is provided as an ogive form having a forward tip displaced at an angle &bgr; above the body longitudinal centerline

38

. The nose angle &bgr; is fixed when the nose assembly

16

is disposed on the body

12

. A plurality of designs for the nose assembly

16

can be used, providing the nose angle &bgr; ranging from approximately 0° to approximately 12° above the body longitudinal centerline

38

. The purpose for the nose angle &bgr; is to provide upward thrust by creating an angle-of-attack at the forward end of the body

12

. The additional lift provided by the nose angle &bgr; helps to compensate for the aft location of the arch wing

14

which is provided aft of the RLV

10

center of gravity

45

. The steering device

18

includes two edges

47

formed above and below the body longitudinal centerline

38

. Each of the edges

47

are disposed at an angle from the body longitudinal centerline

38

of approximately 10° to approximately 20°. A joint at the forward end of each of the steering devices

18

is attached to the body

12

such that the steering devices

18

rotate about a rotation axis

46

.

FIG. 4

also shows in phantom a payload discharge section D and a payload stowage section E. In a preferred embodiment, the payload stowed in the payload stowage section E is displaceable through the payload discharge face

36

. To provide for an aft discharge from the RLV

10

, a portion of the payload stowage section E can overlap the payload discharge section D providing clearance for the payload to pass within and through the payload discharge section D. Additional components (not shown) are stowed in the RLV

10

including navigation equipment, power sources, and additional control and operational systems. These components are common to reusable platforms or weapon delivery vehicles and are thus not shown for clarity.

Referring to

FIG. 5

, the RLV

10

is shown during a re-entry portion of the powerless flight phase. The RLV

10

is operable at hypersonic speeds, therefore following launch, the RLV

10

is designed to dissipate the heat of re-entry at lowered hypersonic speeds by re-entering at a high angle-of-attack &mgr; as shown. The angle-of-attack &mgr; ranges between approximately 30° to approximately 60°. A thermal insulation layer

48

is disposed on an underside of the body

12

. A similar wing thermal insulation

50

is disposed on an under- surface of the lower wing

22

. The thermal insulation layer

48

and the wing thermal insulation layer

50

are preferably silicon-based materials similar to those used on reusable platforms known in the art. The wing thermal insulation layer

50

radiates heat outwardly when passing through a hypersonic flow field as shown. By displacing the lower wing

22

forward of the upper wing

20

, when the RLV

10

is operated at the angle-of-attack &mgr; for re-entry, the lower wing

22

provides a shielded flow region F wherein hypersonic flow generated temperatures are avoided due to shielding of the upper wing

20

. Thermal insulation material is therefore not required on the upper wing

20

. The high angle-of-attack &mgr; is used at reentry velocities above approximately Mach 10. At velocities below approximately Mach 10, the high angle-of-attack &mgr; is not required. A horizontal flight path is indicated by arrow G.

Referring to

FIG. 6

, another preferred embodiment of a reusable platform

100

includes a body

102

having a ring canard

104

rotatably attached in place of the steering device

18

(shown in FIG.

1

). The ring canard

104

is disposed about a canard axis-of-rotation

106

. Each of the starboard and port sides of the ring canard

104

can be actuated in unison or independently of each other. By rotating the ring canard

104

about the canard axis-of-rotation

106

, the ring canard

104

deflects as shown in the exemplary forward rotation direction H and the aft rotation direction J. By selectively rotating the ring canard

104

, forward attitude control of the reusable platform

100

is obtained.

In a preferred embodiment, the ring canard

104

is initially stowed during the launch phase of operation. The stowed position (not shown) provides the ring canard

104

wrapped generally about the perimeter of the body

102

. A mechanism (not shown) releases the ring canard and a spring tension inherent in the design displaces the ring canard

104

into its operating position shown. Materials for the ring canard

104

can include spring steel and composite materials capable of producing the spring force necessary to position and hold the ring canard

104

in its operating position.

FIG. 6

also shows a plurality of exemplary designs for a nose assembly. A 0° ogive nose

108

is shown installed on the body

102

. A positive 5° ogive nose

110

and a positive 10° ogive nose

112

are also shown. Accordingly, a plurality of individual nose assemblies can be installed on the body

102

to affect the lift force generated by the nose assembly.

Referring to

FIG. 7

, an elevation view showing the reusable platform

100

of

FIG. 6

is provided. The ring canard

104

is shown in each of 3 potential operating positions when rotated about the canard axis of rotation

106

(shown in FIG.

6

). In a forward rotated position a download is applied to the reusable platform

100

via the ring canard

104

. In a neutral or centrally vertical position, the ring canard

104

will produce additional upload as shown if the reusable platform

100

is operating at or above an operating angle-of-attack &dgr; as shown. If the reusable platform

100

is operating at a horizontal flight path indicated by horizontal flight path direction arrow G, a central or perpendicular position of the ring canard

104

will generate no load on the reusable platform

100

. In the horizontal flight path G, an aft positioned ring canard

104

will produce an upload for the reusable platform

100

. The advantage of the ring canard

104

is the capability of producing either an upload or a download depending upon the rotated position. The upload or download force is applied at the forward end of the reusable platform

100

and acts to rotate the reusable platform

100

about the platform center of gravity

114

. This provides pitch control for the reusable platform

100

. By rotating individual sides of the ring canard

104

in the forward rotation angle H and the aft rotation angle J (shown in FIG.

6

), side force can be generated to turn the reusable platform

100

starboard or port.

As best seen in

FIG. 8

, in a preferred embodiment, the RLV

10

of the present invention discharges payload (e.g., sub-munitions) in a rear-release direction. During the weapons release phase, a payload

52

is released from the aft end of the RLV

10

in the payload discharge direction A. A plurality of known means can be used to discharge the payload

52

. Exemplary means for discharge include using a drag parachute (not shown) to pull the payload

52

from its stowed position, or an ejection device (not shown) to propel the payload

52

in an aft direction, triggered either by an onboard signal or a remote signal received by the RLV

10

. In the exemplary application of sub-munitions shown in

FIG. 8

, a sub-munition guidance system

54

as known in the art is employed for self propulsion of the payload

52

to a target or ground site. The RLV

10

can carry a plurality of payloads, and can deliver sub-payloads to multiple locations.

Referring to

FIG. 9

, the steering device

18

shown in

FIG. 1

is provided in greater detail. The steering device

18

, when deployed in its steering device deployed position C (shown in FIG.

2

), produces a lateral force which directs the forward end of the RLV

10

in the desired direction. At a steering device deployment angle &agr;, boundary layer air flow propagates through the air flow clearance

42

to maximize the effectiveness of the steering device

18

. The steering device

18

rotates about a rotation axis

46

to achieve any steering device deployment angle &agr; between approximately 0° to approximately 30°. Material for the steering device is preferably selected from the group of low dielectric materials able to reduce radar reflection.

Referring to

FIG. 10

, a cross section taken through the arch wing

14

and the body

12

shows exemplary material layers and an exemplary grouping of submunitions. The wing thermal insulation layer

50

is preferably a silica-based material to reflect reentry induced heating at the lower wing

22

. The lower wing

22

can be constructed of metal, including titanium or aluminum, or a composite material, and preferably includes a radar absorbing material (RAM) layer

68

preferably coated on both surfaces of the lower wing

22

, but at least those surfaces of the lower wing

22

not covered by the thermal insulation layer

50

. The lower wing

22

is bonded (e.g., by adhesive) or mechanically joined (e.g., by rivets) to the curved end plates

24

at lower wing joints

70

and

72

, respectively. The curved plates

24

are in turn bonded or mechanically joined to the upper wing

20

at upper wing joints

74

and

76

, respectively. Because no metal structural members are required in their cross sections, the material for the curved plates

24

and the upper wing

20

is preferably a radar translucent material such as a low dielectric material.

FIG. 10

also shows that a RAM layer

78

is applied to the outer surface of the body

12

. An exemplary payload cylinder

80

is shown in its installed position, having a plurality of sub-munition chambers

82

. A stowage cylinder

84

is also shown, for stowing an inflatable tail cone and extendable shaft (items

404

and

406

, respectively, shown and discussed in reference to FIG.

11

).

As best seen in

FIG. 11

, another preferred embodiment includes a reusable platform

400

having a body

402

. An arch wing is not shown in the details of

FIG. 10

for clarity. Following the powerless glide phase, the reusable platform

400

begins the landing phase of operation. During the landing phase it is important for the reusable platform

400

to obtain a maximum free glide distance to maximize the RLV operating range. To minimize aerodynamic drag, and following the release of the payload, the reusable platform

400

further includes an inflatable tail cone

404

. The inflatable tail cone

404

extends via an extendable shaft

406

from an aft end

408

of the reusable platform

400

. Following extension, the inflatable tail cone

404

self-inflates to form a general cone shape providing clearance to the aft end

408

. The inflatable tail cone

404

includes a rounded leading face

410

and a tapering body

412

. Following the landing phase the inflatable tail cone

404

and the extendable shaft

406

are reloaded through the aft end

408

to a stowed position

414

(shown in phantom).

Referring now to

FIG. 12

, the plurality of components installed on the lower wing

22

are shown. A starboard elevon

56

, a port elevon

58

, a starboard aileron

60

, and a port aileron

62

are installed adjacent to an aft edge

64

of an upper surface

66

of the lower wing

22

. The starboard elevon

56

and the port elevon

58

are installed inboard of the starboard aileron

60

and the port aileron

62

. The positions of the elevons and the ailerons can be modified from those shown in

FIG. 12

depending upon the geometry of the lower wing

22

and the flight characteristics of the RLV

10

. As previously noted, the curved end plates

24

prevent higher pressure air under the lower wing

22

from bypassing the flight surfaces of the lower wing

22

by escaping to the upper surface

66

of the lower wing

22

. Air flow is therefore maintained over the upper surface

66

in the general direction of flow direction arrows K.

The RLV of the present invention offers several advantages. The arch wing design increases the available lift by approximately 50% over a single planar wing. The end plates joining both the upper wing and the lower wing of the arch wing provide directional stability to the air vehicle and maintains air flow over the flight control surfaces of the lower wing in a flow direction generally perpendicular with the surface of the lower wing. The nose assembly design provides additional lift through the use of a fixed angle-of-attack. The forward steering devices are located forward of the center of gravity of the reusable platform therefore generating side thrust and enabling additional pitch control for the RLV. The slender and rounded surfaces of the reusable platform including the swept arch wing reduce the radar cross section.

The capability of achieving hypersonic speeds by boost from a separate launch platform or detachable booster rockets enables the RLV to reach targets at higher than sonic velocities, reducing the likelihood of enemy detection and destruction of the RLV. In the preferred embodiment, by releasing the payload in an aft direction, payload doors which generate aerodynamic drag and radar reflective cavities are not required, therefore the RLV is able to maintain higher than sonic velocities during the powerless flight and weapons release phases. Radar absorbing or radar translucent materials are provided at exterior surfaces of the body, the forward steering device, and for wing portions of the arch wing. A reusable platform of the present invention can be used at both atmospheric and low earth orbital elevation, increasing the range of operation. The arch wing of the present invention in concert with the forward steering devices provide sufficient vertical stabilization such that a rudder, winglets, or other vertical stabilizers known in the art are not required for control of the RLV.

The materials selected for use for the RLV can vary depending upon the application and temperature range of operation. Composite materials can be used for the arch wing provided that the thermal insulation protects the lower temperature capability areas of the arch wing. Higher strength materials including titanium and steels can also be used for the body, the arch wing, the forward steering devices, and the nose assembly of the reusable platform. An RLV of the present invention is capable of velocities up to approximately Mach 30 when provided with thermal insulation materials on selected under-surfaces to protect the under-surfaces during an RLV re-entry phase. The powerless gliding phase is initiated at above Mach speed velocities providing increased range of operation with reduced chance of detection.

The description of the invention is merely exemplary in nature and, thus, variations that do not depart from the gist of the invention are intended to be within the scope of the invention. Such variations are not to be regarded as a departure from the spirit and scope of the invention. For example, a rocket engine is described as the means for achieving launch for the reusable platform of the present invention, however, the RLV can also be launched by other means. The launch phase is performed by a separate craft known in the art carrying the RLV to a desired elevation and/or by booster rockets temporarily attached to the body of the RLV as known in the art.