Method and apparatus for control of shock/boundary-layer interactions

This application claims the benefit of U.S. Provisional Application No. 60/297,568, filed Jun. 12, 2001.

STATEMENT OF GOVERNMENT INTEREST

This invention was made with Government support under Contract No. F49620-98-1-0490 awarded by the U.S. Air Force Office of Scientific Research (AFOSR). The Government has certain rights in this invention.

FIELD OF THE INVENTION

The present invention relates generally to control of shock/boundary-layer interactions caused by supersonic airflow, and more particularly to configurations and actuation of flaps used in control of shock/boundary-layer interactions.

BACKGROUND OF THE INVENTION

Shockwaves are encountered when an aircraft reaches supersonic airspeeds. Such shockwaves exert significant forces on the thin layer of air around the aircraft, a component referred to as the boundary layer. These shockwaves interact with the boundary layer and, during strong interactions, can cause the boundary layer to be degraded, and may also induce high levels of flow separation. These undesired boundary layer interactions accordingly bring about safety, performance, and longevity concerns, especially when the interactions occur inside of engine inlets.

Systems for alleviating such interactions have been developed. These systems bleed air near the boundary layer to suppress shockwave induced flow separation and improve overall flow uniformity. Active transpiration systems require some sort of ducting and/or pumping to bleed the air, which occupies valuable space, and increases the overall weight and cost of the vehicle.

One alternative to boundary layer bleed is to use cavity recirculation. This passive transpiration control method consists of a porous surface and a cavity underneath. The porous surface can be made of holes or slots. During supersonic flight, the changes in pressure will cause air downstream of the shock impingement to flow into the holes, through the cavity and then out through the holes upstream of the impingement. These systems have reduced mechanical complexity and expense compared to the conventional active transpiration systems. However, present models for passive transpiration systems have disadvantages. Transpiration rates are typically insufficient for effective boundary layer control due to the hole aerodynamics. For example, holes or slots that are normal to the surface create a geometry that is significantly less effective than angled holes for bleeding purposes. Further, the holes can yield increased drag at lower Mach speeds or subsonic air flight because of their continuous open state. This potential leads to the same design concerns experienced in needing to determine the location of shock boundary interaction in a particular aircraft so the holes can be limited to that area. Otherwise, drag losses become too significant.

SUMMARY OF THE INVENTION

The present invention is directed to an apparatus for controlling shock/boundary-layer interactions created by a supersonic shock on a surface of a structure. The apparatus includes a cavity formed in the structure and having an opening on the surface. A plate is attached to the surface and covers the opening. A plurality of flaps are formed on the plate and is operable to cooperatively close the opening in response to subsonic airflow condition over the flaps, and open the opening to permit airflow through the cavity in response to supersonic airflow conditions over the flaps.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1A

is a diagram of a system for controlling shock/boundary-layer interaction in accordance with the present invention in a condition responsive to subsonic airflow;

FIGS. 1B and 1C

are diagrams of the system for controlling shock/boundary-layer interaction in accordance with the present invention in conditions responsive to oblique and normal shocks, respectively, caused by supersonic airflow;

FIG. 2

is a top view of one embodiment of the flaps of the control system of the present invention;

FIG. 3

is a sectional view of a stringer plate on which the flaps of

FIG. 2

are adapted to be mounted;

FIG. 4

is a top view of another embodiment of the flaps of the control system of the present invention;

FIG. 5

is a top view of yet another embodiment of the flaps of the control system in accordance with the present invention;

FIG. 6

is a top view of further embodiment of the flaps of the control system in accordance with the present invention;

FIG. 7

is a top view of one embodiment of the flaps of the control system in accordance with the present invention in which upstream flaps are replaced with a plurality of holes;

FIG. 8

is a graph illustrating the affect of heat on nitinol;

FIG. 9

is a diagram showing heaters attached to flaps for controlling flap deflection in accordance with another embodiment of the control system of present invention; and

FIG. 10

is a block diagram of the control system of the present invention incorporating the heaters of FIG.

9

.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to

FIG. 1A

, the operation of a shock/boundary-layer interaction (SBLI) control system

10

in accordance with one embodiment of the invention is illustrated in a condition responsive to subsonic airflow. The control system

10

includes a plurality of injection flaps

12

and bleed flaps

14

which control airflow through a cavity

16

bounded by physical barriers

17

on its remaining sides. During subsonic airflow, the mesoflaps

12

,

14

remain closed over the cavity

16

. In this condition, the cavity

16

and flaps

12

,

14

create no interruption of a boundary layer

18

because the flaps effectively inhibit air circulation through the cavity

16

. Thus, the subsonic flow condition of the flaps

12

,

14

is an effectively smooth surface over which boundary layer

18

passes freely and without interference or added drag. The closed condition of the flaps is maintained under subsonic airflow as a result of the uniform or nearly uniform pressure between the boundary layer

18

and the cavity

16

. No-shock conditions indicative of subsonic flow create a nearly constant streamwise pressure distribution on the flaps

12

,

14

. Thus, the pressure in the cavity

16

is nearly equal to that in the flow above the flaps

12

,

14

.

As is known in the art, the transition to supersonic airflow is accompanied by an oblique shock

20

, as illustrated in

FIG. 1B

, or by a normal shock as shown in FIG.

1

C. Such shocks

20

create strong streamwise pressure variations, with an area of high pressure downstream of the shock

20

and an area of low pressure upstream of the shock

20

. The pressure variation created during supersonic airflow conditions deflects the flaps

12

,

14

and creates the desired air circulation through the cavity

16

. High downstream pressure causes the injection flaps

12

to deflect into the cavity

16

and direct the boundary layer

18

into the cavity

16

. The nearly constant pressure in the cavity

16

will lie roughly between the high downstream pressure and low upstream pressure. Thus, the bleed flaps

14

upstream of the shock

20

deflect out of the cavity

16

to direct the airflow out of the cavity, thus circulating the boundary layer

18

as shown in FIG.

1

B.

By circulating the boundary layer

18

, the control system

10

reduces the interaction between the boundary layer and the shocks

20

that occurs at supersonic airflow conditions. The flaps

12

,

14

direct the boundary layer

18

into and out of the cavity

16

at an angle. Angled active bleed systems have been previously shown to be more effective in controlling the interaction between the shock

20

and the boundary layer

18

than passive systems that direct perpendicular airflow.

Turning now to

FIG. 2

, the flaps

12

,

14

are formed on a flap plate

22

, and include a portion

24

extending generally transversely to the direction of the boundary layer

18

. A longitudinal portion

26

extends from each end of the transverse portion

24

generally perpendicularly to the transverse portion, and in the direction opposite to that of the boundary layer

18

. Each of the flaps

12

,

14

have rounded corners

28

in order to minimize corner flap flutter and peak stress locations and a hole

30

at the distal end of each longitudinal portion

26

to relieve stress to the flaps as they are deflected.

In one embodiment, the flap plate

22

is fabricated from a nickel-titanium alloy known as nitinol, and includes four flaps

12

,

14

. It should be understood that each of the flaps can be injection flaps

12

or bleed flaps

14

depending on the location of the shock

20

. The length of the transverse portion

24

of the flaps

12

,

14

can be anywhere from approximately 1 to 20 boundary layer thicknesses, but is preferably about 10 boundary layer thicknesses. The length of the longitudinal portions

26

can be anywhere from approximately 1 to 3 boundary layer thicknesses. The thickness of the flaps

12

,

14

is preferably less than ¼ of a boundary layer thickness. The flaps

12

,

14

are approximately 1 boundary layer thickness apart from each other.

It should be noted that the width and the length of the cavity

16

generally correspond with those of the flap plate

22

. The cavity depth should be at least one boundary layer thickness deep to provide adequate recirculation of the boundary layer

18

.

Referring to

FIG. 3

, the flap plate

22

is configured and adapted to be mounted to a stringer plate

32

, which supports the flaps

12

,

14

in all static areas. The stringer plate

32

includes a plurality of openings

34

corresponding to each of the flaps

12

,

14

, and configured to allow the flaps to deflect up or down. Spars

36

formed adjacent each opening

34

and extending the width of the stringer plate

32

, are machined to a sharp angle for aerodynamic purposes. In the preferred embodiment, the thickness of the stringer plate

32

is approximately less than one boundary layer thickness.

It should be understood that the four-flap configuration of the control system

10

described in

FIG. 2

is only one embodiment of the present invention, and that other configurations are also contemplated. The flap plate

22

may have more or less than four flaps, and the number of flaps may be even or odd. For example, a six-flap configuration in accordance with another embodiment of the present invention is shown in FIG.

4

. In the embodiment shown in

FIG. 4

, the basic shape of the flaps

12

,

14

are the same as those of FIG.

2

. The six-flap configuration, however, is generally thinner than that of the four-flap configuration, though the general dimensions are similar. An increased number of flaps can provide increased performance when the ratio of shock height to boundary layer thickness becomes large.

Turning now to

FIG. 5

, and in accordance with another embodiment of the present invention, the flap plate

22

includes flaps

12

,

14

that have generally “S” shaped longitudinal portions

26

. By curving the longitudinal portions

26

, sharp corners are avoided at both the trailing edge

38

of the flap (which significantly cuts down on flap flutter) and at the upstream edges

40

(which significantly reduces local stress levels and reduces fatigue failure). This configuration of the flaps

12

,

14

also results in the airflow downstream of the curved portion being generally healthier (i.e., higher skin friction) and has stronger performance gains in terms of stagnation pressure recovery. In this context, skin friction is the shear stress the boundary layer

18

exerts on the surface and is generally indicative of “healthier” boundary layers. The stagnation pressure recovery is the integrated total pressure distribution downstream of the interaction and when it is higher, it is generally indicative of a lower gas dynamic drag on the system. The length of the transverse portion

26

of the flaps

12

,

14

can be anywhere from approximately 1 to 3 boundary layer thicknesses, and the length of the longitudinal portions

26

can also be anywhere from approximately 1 to 3 boundary layer thicknesses. The thickness of the flaps

12

,

14

is preferably less than ¼ of a boundary layer thickness, and the flaps are approximately 1 boundary layer thickness apart from each other.

Turning now to

FIG. 6

, and in accordance with another embodiment of the present invention, the injection flaps

12

are separated from the bleed flaps

14

to form two groups of one-way flaps. In other words, the flaps upstream of the shock

20

, i.e., the bleed flaps

14

, are designated to bend only upwards, and the downstream flaps (the injection flaps

12

) designated to bend only downwards. Results from computational fluid dynamics have shown that improvement can be made to the performance of the flaps

12

,

14

. In particular it has been shown computationally that the upstream or bleed flaps

14

should be located several (e.g., 5-10) boundary layer thicknesses upstream of the shock

20

. The downstream or injection flaps

12

are found to be efficient when located both near the shock interaction as well as further downstream, e.g. 2 to 5 thicknesses downstream. It should be understood that while this embodiment is illustrated using the flap configuration of

FIG. 5

, other flap configurations, as shown in

FIGS. 2 and 4

, for example, may also be used.

Turning to

FIG. 7

, and in accordance with yet another embodiment of the present invention, the downstream or injection flaps

12

are designated to bend only downwards, as in the description with respect the embodiment shown in FIG.

6

. However, instead of the upstream or bleed flaps

14

, this embodiment employs a plurality of through holes

42

formed generally normal to the flap plate

22

. The holes

42

have a diameter which is approximately on the order of one boundary layer displacement thickness, i.e., approximately ¼ of the boundary layer thickness, as known in the art. The holes are arranged generally in rows in an area that is approximately the same as an area in which the injection flaps

12

are formed. The number of holes is generally consistent with a porosity of about 5%. Alternatively, the embodiment shown in

FIG. 7

may be implement with through holes

42

which have a diameter that is substantially smaller than the boundary layer displacement thickness, on the order of 5 &mgr;m to 60 &mgr;m, for example.

The flaps

12

,

14

described above deflect open at varying degrees depending on the speed of the airflow. As the pressure rise across the shock rises, the flaps

12

,

14

open to a larger degree, thereby circulating more boundary layer

18

through the cavity

16

. Thus, the control system

10

controls the boundary layer

18

and shock

20

interaction at the higher shock strength levels, while also preventing excess drag during lower Mach speeds. It should also be noted that it is the location of the impinging shockwave which determines which of the flaps are injection flaps

12

and which of the flaps are bleed flaps

14

. Thus, those skilled in the art will appreciate that using flaps constructed according to the invention in any general area in which the shocks

20

are a concern eliminates the need for advance knowledge of exact streamwise shock locations.

Results from experiments and computations have shown that there is an optimum amount of flap deflection for a given shock condition. Optimum deflection may depend on the strength and position of the shock which are known to vary in the high-speed boundary layer control applications, e.g., as altitude and Mach number change. In accordance with an embodiment of the present invention, the stiffness of the flaps

12

,

14

are controlled to obtain optimum deflection. The variation in the stiffness of the flaps

12

,

14

is achieved by controlling the temperature of the flux. There is a transformation of the material from martensite to austenite with an increase in temperature. As shown in

FIG. 8

, which shows the measurements of nitinol flap deflection under temperature variations, the transformation yields more than a two-fold increase in the modulus of Elasticity.

Turning now to

FIG. 9

, micro-integrated circuit heaters

44

are employed to vary the stiffness of the flaps

12

,

14

for the case of nitinol mesoflaps. The heaters

44

are substantially the same size (width and length) as and adhered to the flaps

12

,

14

on the side facing the cavity

16

. In operation, the heat from the heaters

44

causes the flaps

12

,

14

to have an increased stiffness because of the phase transformation at higher temperatures, which in turn causes a reduction in the flap deflection.

Turning now to

FIG. 10

, the heaters

44

are operatively connected to a controller

46

which determines the amount power to be supplied to the heaters

44

for controlling deflection of the flaps

12

,

14

. A pressure sensor

48

is located downstream of the shocks

20

and provides feedback signals to the controller

46

, so that adjustments can be made by the controller to maintain the optimum deflection amid changing conditions. The pressure sensor

48

may also be located in the cavity

16

.

While various embodiments of the present invention have been shown and described, it should be understood that other modifications, substitutions and alternatives are apparent to one of ordinary skill in the art. Such modifications, substitutions and alternatives can be made without departing from the spirit and scope of the invention, which should be determined from the appended claims.

Various features of the invention are set forth in the appended claims.