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Title:
INTEGRATED PULSE DETONATION ENGINE IN A LIFTING SURFACE WITH SUPERCIRCULATION
Document Type and Number:
WIPO Patent Application WO/2007/022315
Kind Code:
A3
Abstract:
An integrated aircraft propulsion system includes a pulse detonation engine integrated with a lifting surface to achieve supercirculation. The exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface and is deflected at least partially around the trailing edge of the lifting surface. The circulation-controlled lifting surface provides augmented lift, reduced drag, and an efficient propulsion system through supercirculation and thrust vector control capability, hi a delta wing, a pulse detonation engine can be used to power one or more control jets, located near the apex or the trailing edge of the wing, for controlling vortex breakdown of the vortices created by the leading edge of the delta wing.

Inventors:
GUTMARK EPHRAIM JEFF (US)
ALLGOOD DANIEL CLAY (US)
Application Number:
PCT/US2006/032057
Publication Date:
June 14, 2007
Filing Date:
August 17, 2006
Export Citation:
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Assignee:
UNIV CINCINNATI (US)
GUTMARK EPHRAIM JEFF (US)
ALLGOOD DANIEL CLAY (US)
International Classes:
B64C21/04; B64D27/00
Foreign References:
US7093794B22006-08-22
US3807663A1974-04-30
US6347509B12002-02-19
US2941751A1960-06-21
US4645140A1987-02-24
Attorney, Agent or Firm:
MANCINO, David, A. et al. (STETTINIUS & HOLLISTER LLP, 425 Walnut Stree, Cincinnati OH, US)
Download PDF:
Claims:

1. An integrated propulsion system comprising: a pulse detonation engine; and a lifting surface; wherein the pulse detonation engine is integrated with the lifting surface such that the exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface.

2. The integrated propulsion system of claim 1 , further comprising: a plenum chamber contained within the lifting surface, wherein the exhaust from the pulse detonation engine enters the plenum chamber prior to being expelled in a rearward direction tangentially to the lifting surface.

3. The integrated propulsion system of claim 2, wherein the lifting surface comprises a curved leading edge, generally flat top and bottom surfaces, and a trailing edge^ wherein the trailing edge of the lifting surface is a curved surface; and wherein the exhaust from the pulse detonation engine is deflected at least partially around the trailing edge of the lifting surface.

4. The integrated propulsion system of claim 3, wherein the trailing edge of the lifting surface has a semicircular cross-section.

5. The integrated propulsion system of claim 3, wherein the pulse detonation engine has a curved exhaust nozzle.

6. The integrated propulsion system of claim 3, wherein the pulse detonation engine has a convergent exhaust nozzle.

7. The integrated propulsion system of claim 3, wherein the pulse detonation engine has a divergent exhaust nozzle.

8. The integrated propulsion system of claim 1 , further comprising:

a controllable baffle positioned downstream of the exhaust nozzle of the pulse detonation engine, the controllable baffle being capable of diverting the exhaust of the pulse detonation engine such that thrust vectoring may be achieved.

9. The integrated propulsion system of claim 1 , further comprising: a trailing edge flap attached to the lifting surface, the flap being capable of varying the flow of air over the trailing edge of the lifting surface.

10. The integrated propulsion system of claim 1, further comprising: an augmentor wing joined to the lifting surface and oriented substantially parallel to the lifting surface; wherein the exhaust from the pulse detonation engine enhances the flow of air between the lifting surface and the augmentor wing; whereby increased lift is achieved.

11. The integrated propulsion system of claim 1 , wherein the lifting surface is a delta wing.

12. The integrated propulsion system of claim 11, further comprising: at least one control jet powered by the exhaust from the pulse detonation engine for controlling vortex breakdown of the vortices created by the leading edge of the delta wing.

13. The integrated propulsion system of claim 12, wherein the at least one control jet is located near the apex of the delta wing.

14. The integrated propulsion system of claim 12, wherein the at least one control jet is located near the trailing edge of the delta wing.

15. An integrated propulsion system comprising: a plurality of pulse detonation engines; and a lifting surface;

wherein the plurality of pulse detonation engines are integrated with the lifting surface such that the exhaust from the plurality of pulse detonation engines is expelled in a rearward direction tangentially to the lifting surface.

16. The integrated propulsion system of claim 15, further comprising: a controllable tab positioned downstream of the exhaust nozzles of the plurality of pulse detonation engines, the controllable tab being capable of diverting the exhaust of the plurality of pulse detonation engines such that thrust vectoring may be achieved.

17. The integrated propulsion system of claim 15, further comprising: a trailing edge flap attached to the lifting surface, the flap being capable of varying the flow of air over the trailing edge of the lifting surface.

18. The integrated propulsion system of claim 15, further comprising: an augmentor wing joined to the lifting surface and oriented substantially parallel to the lifting surface; wherein the exhaust from the plurality of pulse detonation engines enhances the flow of air between the lifting surface and the augmentor wing; whereby increased lift is achieved.

19. An integrated propulsion system comprising: a pulse detonation engine; and an ejector having at least two Coanda nozzles; wherein the pulse detonation engine is integrated with the ejector such that the exhaust from the pulse detonation engine is expelled between the Coanda nozzles; and wherein each of the Coanda nozzles may be rotated such that thrust vectoring may be achieved.

20. The integrated propulsion system of claim 19, wherein the pulse detonation engine is integrated with the ejector in such a way that ambient air may flow through a secondary flow channel between the Coanda nozzles and the exhaust outlet of the pulse detonation engine.

21. An aircraft comprising: an airframe; at least one pulse detonation engine; and a lifting surface coupled to the airframe; wherein the pulse detonation engine is integrated with the lifting surface such that the exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface.

22. An aircraft comprising: an airframe; a lifting surface coupled to the airframe; an ejector having at least two Coanda nozzles; and at least one pulse detonation engine integrated with the ejector; wherein the at least one pulse detonation engine is integrated with the ejector such that the exhaust from the at least one pulse detonation engine is expelled between the Coanda nozzles; and wherein each of the Coanda nozzles may be rotated such that thrust vectoring may be achieved.

Description:

INTEGRATED PULSE DETONATION ENGINE IN A LIFTING SURFACE WITH SUPERCIRCULATION

CROSS-REFERENCE TO RELATED APPLICATION , [0001] This application claims the benefit of U.S. Provisional Patent Application Serial No. 60/709,194, filed August 18, 2005, and U.S. Non-provisional Patent Application Serial No. 11/413,981, filed April 28, 2006, the disclosures of which are incorporated herein by reference.

BACKGROUND

[0002] The present invention relates to an aircraft propulsion system using a pulse detonation engine integrated with a lifting surface to achieve combined benefit of augmented lift, reduced drag, and an efficient propulsion system through supercirculation and thrust vector control capability.

[0003] The benefits of propulsion-airframe integration have been primarily demonstrated at hypersonic speeds. At these speeds, a scramjet engine is completely integrated with the airframe. At supersonic speeds, it is possible to utilize the flow fields induced by the engine nacelles to provide favorable interference drag reductions and interference lift. Conversely, the airframe (body or wings) can be used to precompress the flow entering the engineinlets for improved engine performance. At subsonic speeds, such integration research and analysis was only used to alleviate problems or unfavorable interactions. Exploiting propulsion-airframe integration to lower speeds may lead to more efficient aircraft by producing higher lift to drag ratio, better cruise efficiency, increased maneuverability and ultimately lead to new vehicle designs.

SUMMARY

[0004] An integrated aircraft propulsion system includes a pulse detonation engine integrated with a lifting surface to achieve supercirculation. The exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface and is deflected at least partially around the trailing edge of the lifting surface. The circulation-controlled lifting surface provides augmented lift, reduced drag, and an efficient propulsion system through supercirculation and thrust vector control capability.

[0005] Accordingly, it is a first aspect of the present invention to provide an integrated propulsion system including: a pulse detonation engine and a lifting surface, where the pulse detonation engine is integrated with the lifting surface such that the exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface. In a detailed embodiment, the trailing edge of the lifting surface is a curved surface, and the exhaust from the pulse detonation engine is deflected at least partially around the trailing edge of the lifting surface. In various embodiments, the pulse detonation engine can have a curved, convergent, or divergent.

[0006] hi an alternative detailed embodiment of the first aspect of the present invention, the integrated propulsion system further includes a controllable baffle positioned downstream of the exhaust nozzle of the pulse detonation engine and capable of diverting the exhaust of the pulse detonation engine such that thrust vectoring may be achieved, hi another alternative detailed embodiment, a trailing edge flap attached to the lifting surface, the flap being capable of varying the flow of air over the trailing edge of the lifting surface. In another alternative detailed embodiment, an augmentor wing can be joined to the lifting surface and oriented substantially parallel to the lifting surface, where the exhaust from the pulse detonation engine enhances the flow of air between the lifting surface and the augmentor wing; thus achieving increased lift.

[0007] hi an alternative detailed embodiment of the first aspect of the present invention, the lifting surface can be a delta wing, hi this embodiment, at least one control jet powered by the exhaust from the pulse detonation engine can be provided, either near the apex or the trailing edge of the wing, for controlling vortex breakdown of the vortices created by the leading edge of the delta wing.

[0008] It is a second aspect of the present invention to provide an integrated propulsion system including: a plurality of pulse detonation engines and a lifting surface, where the plurality of pulse detonation engines are integrated with the lifting surface such that the exhaust from the plurality of pulse detonation engines is expelled in a rearward direction tangentially to the lifting surface. In a detailed embodiment, a

controllable tab can be positioned downstream of the exhaust nozzles of the plurality of pulse detonation engines, the controllable tab being capable of diverting the exhaust of the pulse detonation engines such that thrust vectoring may be achieved. Additionally, the second aspect of the present invention may be practiced with any of the features or embodiments, or any combination thereof, described above with reference to the first aspect.

[0009] It is a third aspect of the present invention to provide an integrated propulsion system including: a pulse detonation engine and an ejector having at least two Coanda nozzles; where the pulse detonation engine is integrated with the ejector such that the exhaust from the pulse detonation engine is expelled between the Coanda nozzles; and where each of the Coanda nozzles may be rotated such that thrust vectoring may be achieved. In a detailed embodiment, the pulse detonation engine can be integrated with the ejector in such a way that ambient air may flow through a secondary flow channel between the Coanda nozzles and the exhaust outlet of the pulse detonation engine.

[0010] It is a fourth aspect of the present invention to provide an aircraft including: an airframe; at least one pulse detonation engine; and a lifting surface coupled to the airframe; where the pulse detonation engine is integrated with the lifting surface such that the exhaust from the pulse detonation engine is expelled in a rearward direction tangentially to the lifting surface. The fourth aspect of the present invention may be practiced with any of the features or embodiments, or any combination thereof, described above with reference to the first, second, and third aspects.

[0011] It is a fifth aspect of the present invention to provide an aircraft including: an airframe; a lifting surface coupled to the airframe; an ejector having at least two Coanda nozzles; and at least one pulse detonation engine integrated with the ejector; where the at least one pulse detonation engine is integrated with the ejector such that the exhaust from the at least one pulse detonation engine is expelled between the Coanda nozzles; and where each of the Coanda nozzles may be rotated such that thrust vectoring may be achieved. The pulse detonation engine can be integrated with the ejector in such a way that ambient air may flow through a secondary flow channel between the Coanda nozzles and the exhaust outlet of the pulse detonation engine.

[0012] These and other aspects and embodiments will be apparent from the following description, the accompanying drawings, and the appended claims.

BRIEF DESCRIPTION OF THE DRAWINGS

[0013] FIG.1 shows a cross-sectional view of an exemplary embodiment of a circulation-controlled lifting surface.

[0014] FIG.2 shows a cross-sectional view of an exemplary embodiment of a pulse detonation engine.

[0015] FIGS.3 through 6 depict several design features that can be incorporated into the pulse detonation engine to achieve a deflagration-to-detonation transition in the ignited fuel/oxidizer mixture.

[0016] FIG.7 shows a cross-sectional view of the integrated assembly used for experimental testing, including the pulse detonation engine and the lifting surface, according to an exemplary embodiment of the present invention.

[0017] FIGS.8 and 9 illustrate the use of an upper extension piece to the pulse detonation engine exhaust to direct the exhaust flow further around the Coanda surface, according to an exemplary embodiment of the present invention.

[0018] FIGS.10 and 11 illustrate the use a protrusion or step near the exhaust slot to cancel the Coanda effect and allow the exhaust gas to flow in a rearward direction, according to an exemplary embodiment of the present invention.

[0019] FIGS.12 and 13 show a top plan view and a perspective view, respectively, of a lifting surface with a plurality of pulse detonation engines integrated therewith, according to an exemplary embodiment of the present invention.

[0020] FIG.14 depicts an aircraft whose wings are lifting surfaces into which pulse detonation engines have been integrated, according to an exemplary embodiment of the present invention.

[0021] FIG.15 shows a cross-sectional view of an integrated assembly including a pulse detonation engine, a lifting surface, and a trailing edge flap, according to an exemplary embodiment of the present invention.

[0022] FIG.16 shows a cross-sectional view of an exemplary embodiment of an augmentor wing.

[0023] FIG.17 shows an ejector that uses Coanda surfaces to augment the trust from a PDE, according to an exemplary embodiment of the present invention.

[0024] FIGS .18 through 20 are schematic diagrams illustrating the primary leading edge vortices on a delta wing and the use of apex control jets to delay vortex breakdown, according to an exemplary embodiment of the present invention.

[0025] FIG.21 shows a plurality of pulse detonation engines integrated with a delta wing, where some of the plenum chamber exhaust slots have been diverted to exhaust in a rearward direction, thus aiding in the delay of vortex breakdown, according to an exemplary embodiment of the present invention.

DETAILED DESCRIPTION

[0026] The present invention addresses the integration of a pulse detonation engine (PDE) into a lifting surface in order to achieve combined benefit of augmented lift, reduced drag, and an efficient propulsion system through supercirculation (also called powered lift) and thrust vector control capability. The terms "supercirculation" and "powered lift" refer to use of propulsion bleed and/or exhaust flows to increase airflow over a lifting surface. The addition of such a powered blowing mechanism to the wing results in a greater rate of airflow over the wing than would be achieved solely from the ambient airflow rate, thus increasing the lift produced by the wing. This is accomplished through various means: entrainment of external flows, super- velocity acceleration of flows, and direct vectoring of propulsive flows in the lift vector orientation.

[0027] An example of the first concept is the circulation-controlled wing. Circulation control refers to the increase of air circulation over a lifting surface or wing by means of blowing additional air over the wing in a rearward direction, tangential to the wing. On a typical lifting surface, the flow from the upper surface cannot turn around the sharp trailing edge and therefore the flow separates from the trailing edge and the rear stagnation line becomes fixed to the trailing edge of the wing (Kutta condition). For a given angle of attack, separation at the trailing edge occurs for a particular value of the circulation and, hence, for a particular lift coefficient. In order to achieve greater circulation, a lifting surface with a curved trailing edge may be used.

[0028] A schematic diagram of an exemplary circulation control lifting surface 10 is shown in FIG.l. As shown, the circulation control lifting surface 10 has a rounded trailing edge 12. When air flows over the top surface of this lifting surface and reaches the trailing edge 12, it tends to remain attached to the trailing edge 12 and is therefore deflected downward. This tendency of a fluid stream to remain attached to a convex surface rather than follow a straight line in its original direction is called the "Coanda effect," and a surface that causes such a deflection of a fluid stream may be called a "Coanda surface."

[0029] Without powered blowing, the circulation control lifting surface 10 will have a separation point Sl that is partway around the curve of the trailing edge 12 (the Coanda surface). The precise location of this separation point, and hence the amount of deflection produced by the Coanda surface, will be determined by the speed of the airflow over the lifting surface. By increasing the speed of the fluid stream with powered blowing, the separation point can be moved further around the trailing edge to a point S2. The blowing can be accomplished by expelling exhaust gases from a chamber inside the wing through slot 14 near the trailing edge such that the exhaust flow from the slot is tangent to the lifting surface. The slot flow is at a higher speed than that of the local outer-flow and thus energizes the mixing boundary 16. This action permits the upper flow over the lifting surface to remain attached until it reaches the separation point S2. From inviscid theory, the separation point S3 for the boundary layer on the wing's lower surface coincides with S2; however, for a viscous fluid a "dead air" region can exist, with S2 and S3 at its extremities. The interaction

between the outer inviscid flow and the jet flow shifts the rear stagnation point towards point S3, consequently, the circulation is increased and so is the wing's lift.

[0030] The lift of a circulation control lifting surface is a direct function of turbulent mixing between the upper surface boundary layer 18 and the wall jet issued from the exhaust slot 14. This mixing mechanism determines the limit of circulation increase before the flow separates again from the upper surface. In the present invention, the mechanism of periodic excitation of the injected flow is inherent to the system due to the pulsatile nature of the PDE operation.

[0031] The powered lift concepts described herein achieve increased lift by increasing wing circulation and by deflecting thrust downward. These systems can significantly increase the maximum lift coefficient (C^ max ) of the aircraft and thus provide short takeoff and landing (STOL) capability, hi the clean configuration, i.e., without any flaps extended, wings are limited to C^ max values well below 1.5. Mechanical flaps can increase C^ max to around 2.0. Blowing boundary layer control (BLC) is limited to values around 4. For C^ max values above 4, forced circulation is required; further increases require the addition of direct thrust. Application of supercirculation during flight can also help to reduce the angle of attack and induced drag. In addition to supercirculation, the integration of the propulsion system with the airframe can provide the combined benefit of augmented lift, reduced drag, increased maneuverability, and an efficient propulsion system through supercirculation and thrust vector control capability.

[0032] In order to achieve the supercirculation or powered lift described above, a source for the powered blowing of air over the Coanda surface to increase the speed of the airflow over the Coanda surface can be provided. The present invention integrates a pulse detonation engine with a wing in order to generate such powered lift. The PDE is a strong candidate for integrated propulsion-airframe systems due to its simplicity and flexibility in geometry and ease in scalability anywhere from miniature devices to large-scale engines. Recent research on PDEs has revealed their potential as an efficient, low cost propulsion system. The PDE cycle offers simplicity with few moving parts, high thrust to weight ratios, low cost, and ease of scaling. Potentially, the flight operating range of a PDE ranges from static conditions to

hypersonic flight Mach numbers. As of today, there is no other single-cycle propulsion system available with such a broad range of capability. Furthermore, the fact that a PDE does not require a compression cycle eliminates the necessity for heavy and expensive compressor and turbine units, allowing the engines to have very flexible non-axisymmetric geometries. This makes the PDE the ideal candidate for integrated propulsion-airframe systems. Additionally, the pulsating exhaust flow of a PDE will enhance mixing with the ambient air, thus adhering to the surface resulting in higher increased circulation. The following is a non-exhaustive list of references, each of which is incorporated herein by reference, that may be consulted for further information and detail regarding PDEs:

1 Bussing, T. and Pappas, G., "An Introduction to Pulse Detonation Engines," AIAA Paper 94-0263, January 1994.

2 Kailasanath, K. "Recent Developments in the Research on Pulse Detonation Engines", AIAA Journal, Vol. 41, No. 2, February 2003, pp. 145-159.

3 Heiser, W.H., and Pratt, D.T., "Thermodynamic Cycle Analysis of Pulse Detonation Engines," AIAA Journal of Propulsion and Power, Vol. 18, No. 1, 2002, pp. 77-83.

4 Cambier, J. L. and Tegner, J. K., "Strategies for Pulsed Detonation Engine Performance Optimization", Journal of Propulsion and Power, Vol. 14, No. 4, July- August, 1998.

5 Wu, Y., Ma, F. and Yang, V., "System Performance and Thermodynamic Cycle Analysis of Air-breathing Pulse Detonation Engines", Journal of Propulsion and Power, Vol. 19, No. 4, July-August 2003.

6 Allgood, D., Gutmark, E., Rasheed, A. and Dean, A., "Experimental Investigation of a Pulse Detonation Engine with a Two-Dimensional Ejector", AIAA Journal, Vol. 43, No. 2, 2005, pp. 390-398.

7 Schauer, F., Stutrud, J., and Bradley, R., "Detonation Initiation Studies and Performance Results for Pulse Detonation Engines," AIAA Paper 2001-1129, January 2001.

[0033] A schematic of an exemplary embodiment of a PDE 30 is shown in FIG.2. The PDE tube 32 is constructed of stainless steel tubing that is 1 inch in diameter and 24 inches long. The reactants, ethylene and oxygen, are supplied to the PDE 30

through two ports 34. The fuel and oxidizer are injected at a maximum frequency of 50Hz through the use of fast response solenoid valves. The injected mixture is ignited using a standard automotive spark plug 36. In the exemplary embodiment, removable Schelkin spirals 38 are mounted to ring 40, which is inserted between adjoining sections of the PDE tube 32, and are used to accelerate the deflagration-to-detonation process, as explained below. FIG.3 provides a cross-section of the removable Schelkin spirals 38, which fit longitudinally inside the PDE tube 32 and are attached to the mounting ring 40. High frequency response pressure transducers 42 can be mounted along the axis of the PDE tube 32 to monitor the progression of the detonation wave. The PDE exhaust exits from the PDE tube 32 through exhaust nozzle 44.

[0034] In an exemplary embodiment, the exhaust nozzle 44 has a curved shape. In a detailed embodiment, the curved shape can be a convergent shape to help maintain the detonation chamber pressure and extend the blow-down process. This results in increasing or maintaining the thrust of the PDE over a broader range of flight conditions. In an alternative detailed embodiment, a divergent nozzle might be useful for accelerated blowdown rates in order to reach higher PDE operating frequencies. A convergent and divergent nozzle would be helpful to further expand the exhaust products for specific flight operating conditions. Other nozzle shapes can be used to reduce PDE noise.

[0035] In order to achieve detonation of the fuel/oxidizer mixture, a low energy ignition system can be implemented to cause a deflagration to form initially inside the PDE. This deflagration will then be transformed into a detonation through flame acceleration techniques. In order to maximize the PDE propulsion efficiency, the amount of fuel burned during this deflagration-to-detonation transition (DDT) process should be minimized. There are several different techniques for generating DDT, as depicted in FIGS.3 through 6. The primary objective of the DDT device or methodology is to enhance the flame speed to a point where the chemical and gas dynamics become coupled and self-supporting such that detonation event is inevitable.

[0036] One method of accelerating the flame speed is by enhancing the turbulence levels inside the PDE by placing spirals (as depicted in FIG.3) or obstacles (as depicted in FIG.4) in the flow path. Kuo, K., Principles of Combustion, John Wiley, New York, 1986 (incorporated herein by reference). These devices invoke the formation of large-scale structures that stretch the spark-initiated deflagration flame front. The increased flame surface area causes the flame to be accelerated. As the consumption of the fuel/oxidizer mixture is increased rapidly, the energy released from the chemical reactions cause compression waves to form. Due to the increase in gas temperature in the flame zone, the compression waves coalesce into shocks. The shocks then propagate through the fuel/oxidizer mixture heating the flow to the point of self-ignition. These events can take place at several locations within the PDE resulting in several hot-spots or small "explosions." The interaction of these small explosions with the surrounding walls and with each other cause a cascading effect until a detonation wave is developed. The amount of time and distance required for the DDT event is a function of the fuel/oxidizer properties, the level of turbulence, the level of mixing between the fuel and oxidizer and the boundary conditions of the system (or level of confinement). Results by Lee et al suggest that a spiral configuration is more favorable than obstacles due to the lower pressure drop (or drag) induced by the DDT device. Lee, J. H. S., "Dynamic Parameters of Gaseous Detonations," Annual Review of Fluid Mechanics, Vol. 16, 1984, pp. 311-336 (incorporated herein by reference). Severe pressure losses caused by the DDT devices resulted in the average flame speeds to be well below the theoretical CJ detonation velocity. However, it was observed by Lee et al that if the DDT device was successful in accelerating the flame above a critical high-speed deflagration velocity (~ Vi detonation velocity), the deflagration successfully transitioned into a fully developed detonation wave when the flame exited the spiral or obstacle region into the relatively smooth chamber. Lee, J. H. S., Knystautas, R., and Freiman, A., "High Speed Turbulent Deflagrations and Transition to Detonation in H 2 -Air Mixtures," Combustion and Flame, Vol. 56, 1984, pp. 227-239 (incorporated herein by reference). Their results also showed that the detonation wave development can be delayed or hindered by obstacles with too large of a blockage ratio. Meyer et al performed high-speed visualizations of DDT spirals and have also suggested that a PDE with a spiral is more conducive to detonation because the spiral allows helical and transverse modes of the detonation wave to form, where the hot-spots in these

studies were observed to travel along the spiral's helical geometry. Meyer, T. R., Hoke, J. L., Brown, M. S., Gord, J. R., and Schauer, F. R., "Experimental Study of Deflagration-to-Detonation Enhancement Techniques in a H 2 / Air Pulsed-Detonation Engine," AIAA Paper 2002-3720, July 2002 (incorporated herein by reference).

[0037] Other methods of DDT have been suggested as well. For example, cavities (as depicted in FIG.5) were seen by Smirnov et al to facilitate the generation of a detonation wave due to the unsteadiness it invoked in the flame front and shock wave development. Smirnov, N., and Tyurnikov, M., "Experimental Investigation of Deflagration to Detonation Transition in Hydrocarbon-Air Gaseous Mixtures," Combustion and Flame, Vol. 100, 1995, pp. 661-668 (incorporated herein by reference). Properly shaped head-walls of the PDE (as depicted in FIG.6) have been shown by Gelfand et al to "focus" the reflection of shock waves in such a way that either a high-speed deflagration is initiated or a direct-initiation of a detonation is produced. Gelfand, B. E., Khomik, S. V., Bartenev, A. M., Medvedev, S. P., Gronig, H., Olivier, H., "Detonation and Deflagration Initiation at the Focusing of Shock Waves in Combustible Gaseous Mixture," Shock Waves, Vol. 10, 2000, pp. 197-204 (incorporated herein by reference). Another approach to enhancing the DDT process is to increase the level of free radicals in the flow through rapid mixing of combustion gases with the detonable gas as was shown by Knystautas et al and Thomas et al. Knystautas, R., Lee J. H. S. and Moen, L, "Direct Initiation of Spherical Detonation by a Hot Turbulent Gas Jet," 17th Symp. Int. Combustion Proc, 1978, pp. 1235-1245 (incorporated herein by reference); Thomas, G. O. and Jones, A., "Some Observations of the Jet Initiation of Detonation," Combustion and Flame, Vol. 120, 2000, pp. 392- 398 (incorporated herein by reference).

[0038] FIG.7 provides a cross-sectional view of an integrated assembly 50 used for experimental testing, including the PDE 30 and a lifting surface 52, according to an exemplary embodiment of the present invention. In this embodiment, the lifting surface 52 has a cross-section of an oblong circle or ellipse, with a substantially flat top surface 54, a substantially flat bottom surface 56, a curved leading edge 58, and a curved trailing edge 60, which serves as the Coanda surface. The exhaust of the PDE is fed into a plenum chamber 62, which is a chamber in fluid communication with the PDE exhaust nozzle and located inside the wing. The exhaust emerges from the

plenum chamber 62 from a narrow slot 64 on the top of the lifting surface approximate the trailing edge 60. The exhaust slot 64 extends above the trailing edge 60 of the lifting surface by a small amount (0.1 inches in an exemplary embodiment) such that the exhaust gas is directed in a rearward direction, tangential to the lifting surface.

[0039] The distance between the trailing edge 60 and the exhaust slot 64 can be used as a control/design parameter to control the level of supercirculation. It has been found through experiment that the "near optimum" position for the exhaust slot 64would be at the start of the curvature of the trailing edge Coanda surface, but an adjustable upper extension piece could be used to extend the exhaust jet around the Coanda surface. This effectively moves the stagnation point around the trailing edge of the wing and has been shown to be helpful in turning the detonation exhaust products. The adjustable upper extension piece can also be used to control the extent of the exhaust jet deflection and thereby achieve control on the amount of supercirculation and thrust vector control. FIGS.8 and 9 illustrate the effect of adding an upper extension piece 66 to the PDE exhaust slot 64. In FIG.8, the integrated PDE/lifting surface assembly 50 is shown without the upper extension piece. The exhaust exits the plenum chamber 62 through exhaust slot 64 near the top of the trailing edge Coanda surface 60. The exhaust plume 65 follows the trailing edge Coanda surface 60 until the separation point S2. In FIG.9, an upper extension piece

66 has been added to the PDE exhaust slot 64. The curved shape of this extension piece 66 effectively extending the plenum chamber 62 and moving the exhaust slot 64 partway around the trailing edge Coanda surface 60, thus adding a downward component to the PDE exhaust as it exits the plenum chamber 62. The exhaust plume 65 thus follows the trailing edge Coanda surface 60 over a longer distance until the separation point S2, which is now a greater distance around the Coanda surface 60.

[0040] Additionally, the Coanda Effect can be canceled by placing a protrusion or step upstream of the initiation of the Coanda surface. As seen in FIG.10, a protrusion

67 can be placed at the exhaust slot 64, upstream of the Coanda surface 60. With a protrusion having a height of between approximately 0.5 mm and 2 mm, the resulting exhaust plume 65 will follow a straight rearward path, as shown in FIG.10. If the protrusion is made larger than about 2 mm, the resulting exhaust plume can be

deflected upward. In an alternative embodiment shown in FIG.l 1, a rearward-facing step 69 is used in place of the protrusion 67, and a step height of approximately 1 mm will cause the exhaust plume 65 to follow a straight rearward path. The horizontal position of slot exhaust orifice in relation to the protrusion or step, in addition to the slot size, can be used to vary the angle of deflection or the size of the protrusion or step needed to produce a straight exhaust plume.

[0041] FIGS.12 and 13 show a top plan view and a perspective view, respectively, of an integrated assembly 50, illustrating how the PDE and the lifting surface function together. In the embodiment shown, a plurality of PDEs 30 are shown inside the lifting surface 52, each PDE having an exhaust nozzle 44 feeding into a plenum chamber 62. Each plenum chamber 62 directs the exhaust gases to a narrow slot opening 64, which is located near the trailing edge 60 of the lifting surface. As discussed above, the slot 64 is positioned such that the exhaust gas is directed in a rearward direction, tangential to the lifting surface. As the exhaust gas is expelled through the slot 64 in a rearward direction, it is deflected downward as it remains attached to the trailing edge 60 of the lifting surface due to the Coanda effect. As discussed above, the separation point at which the airstream separates from the lifting surface is moved further around the trailing edge 60, thus increasing the lift produced by the lifting surface. Additionally, the downward deflection of the airstream passing over the trailing edge imparts an upward reaction force on the wing, further increasing the lift. The angle of the exhaust stream, and hence the amount of defection, from one or more plenum chambers 62 can be varied by employing upper extension pieces 66, protrusions 67, or steps 69 at the exhaust slot 64, as described above.

[0042] FIG.14 shows a top plan view of an aircraft 80. Each wing 82 of the aircraft 80 is a lifting surface that contains one or more PDEs 30 integrated with the lifting surface in the manner described herein, with the PDE exhaust directed through a plenum chamber 62 and exhaust slot 64. As it exits the exhaust slot 64,the PDE exhaust is directed tangentially over the trailing edge Coanda surface 60 of the lifting surface. The angle of the exhaust stream, and hence the amount of defection, from one or more plenum chambers 62 can be varied by employing upper extension pieces 66, protrusions 67, or steps 69 at the exhaust slot 64, as described above.

[0043] FIG.15 shows an exemplary embodiment of the integrated assembly 70 including a PDE 72 connected to plenum chamber 73, a lifting surface 74, and a trailing edge flap 76. A leading edge slat 77 can also be included. The periodic exhaust of the PDE 72 can delay or prevent separation of the airstream flowing over the flap 76. The trailing edge flap 76 may take various forms that are known to persons skilled in the art. As described above, the exhaust of the PDE exits from the plenum chamber 73 through a slot 78 in a rearward direction tangential to the lifting surface, where the periodic, pulsating nature of the exhaust excites the boundary layer over the trailing edge, which in this embodiment is the flap 76. The frequency and amplitude of the periodic excitation are the key parameters that determine the efficiency of either delaying separation or promoting reattachment of a separated flow from a flap. The optimum frequency at which minimum energy is required for delaying/preventing separation over a flap or promoting reattachment is best described by a normalized frequency F + . F + is defined as the frequency (f) of modulation multiplied by the distance (L f ) between the injection of periodic flow and the trailing edge of the flap, and divided by the external freestream velocity (U∞). More details are given in I. Wygnanski, "Boundary Layer and Flow Control by Periodic Addition of Momentum," Invited talk, AIAA 97-2117, 4 th AIAA Shear Flow Control Conference, June 1997 (incorporated herein by reference).

[0044] Several further application concepts introduce an additional lifting surface to augment the lift induced by the PDE. One such concept is the augmentor wing, shown in FIG.16. As seen in FIG.16, a PDE 72 connected to plenum chamber 73 is integrated with a lifting surface 74. A trailing edge flap 76 and leading edge slat 77 can also be included. Additionally, a second lifting surface 90, called an augmentor wing, is positioned adjacent to the first lifting surface 74. The augmentor wing 90 can have a trailing edge flap 92. The PDE exhaust is expelled from the exhaust slot 78 on the first lifting surface. The resulting exhaust enhances the flow of air between the two lifting surfaces 74 and 90, thus providing increased lift. The trailing edge flap 92 on the augmentor wing 90 can also augment the downward deflection of the PDE exhaust.

[0045] FIG.17 shows another concept, called an ejector, that uses Coanda surfaces to augment the trust from a PDE. In an exemplary embodiment, the engine exhaust

22 is expelled between two Coanda wings 24. Ambient air is drawn through the two secondary flow channels 26 between the Coanda wings 24 and the exhaust nozzle 22, thereby producing added force in the thrust direction. Varying the tilt angle of the two Coanda wings will produce a controllable thrust force that can be vectored. The concept results in a lift force greater than the propulsive force utilized, thus "augmenting" the power output by the engines.

[0046] An additional embodiment is integration of a PDE with a delta wing. In addition to achieving the controlled circulation over the trailing-edge Coanda surface as described above, a PDE integrated with a delta wing can be used to control the location of the breakdown of leading edge vortices inherent in a delta wing. Designers of modern aircraft that require short take-off and landing, aggressive maneuvering in combat, or atmospheric re-entry have to invent new ways of achieving the crucial additional lift. Aircraft flying under these conditions experience large-scale flow separation, or stall, which can result in a loss of lift and control. Aircraft designers have used a variety of different devices including multi-element trailing edge flaps, slats, vortex generators, strakes, and various pneumatic devices in an attempt to control flow separation for high angle of attack flight. Highly swept wings, such as delta wings, address the problems of high angle of attack flight in a different manner. Instead of attempting to prevent flow separation, separation is fixed at the sharp leading edge of the wing and a shear layer is created which rolls up into a pair of strong vortices. As the shear layer rolls up into the primary vortex, it reattaches to the surface and then separates again further outboard on the wing, where a pair of secondary vortices is generated. One is located downstream and outboard of each of the primary vortices Werle, H., Recherche Aeronautique, No 75, pp. 23-30, 1960 (incorporated herein by reference); Earnshaw, P.B. and Lawford, J. A., "Low- Speed Wind-Tunnel Experiments on a Series of Sharp-Edged Delta Wings," Aeronautical Research Council, R& M 3424, 1964 (incorporated herein by reference).

[0047] The primary leading edge vortices are illustrated schematically in FIG.18. In this figure, a full delta wing 100 is shown, including a left-side leading edge 102 and a right-side leading edge 104. Of course, an aircraft fuselage or other vehicle structure could be joined to the wing but is not shown in this figure to avoid over-congestion in the drawing. Using the right side as an example, the primary vortex 110 eminates from the leading edge 104 near the wing's tip 112. The vortex 110 is represented

generally by the swirl 114, indicating the rotation of air. The vortex's direction of rotation is indicated by arrows 116. The core of the vortex, where the strongest axial velocities are found, is represented by the region 118.

[0048] The axial velocity distribution in the primary vortex has a jet-like profile with the velocities in the core being as much as two to three times the freestream velocity. Axial velocity plays an essential role in the existence of leading edge vortices over delta wings. Lee, M. and Ho, C. M., "Lift Force of Delta Wings," Applied Mechanical Review, Vol. 43, No. 9, Sep. 1990, pp. 209-221 (incorporated herein by reference). The continuous flow of air that feeds the vortex at the leading edge is convected away from the wing by the axial velocity. As a result, the vortex is forced to shed from the wing's surface as with typical separated flows. A cross- section of the vortex at 50% of the chord, for example, must allow all of the air that enters the vortex from the entire first half of the wing to pass through it. The highspeed axial flow translates to a low-pressure region over the wing which can account for 30% of the total lift of the wing. Myose, R, Y.., Lee, B. K., Hayashibara, S and Miller, L., S. "An Experimental Study on the Breakdown of Leading-Edge Vortices on Diamond, Cropped, Delta and Double Delta wings During Dynamic Pitching," AIAA Paper # 97-1930, 1997 (incorporated herein by reference).

[0049] A significant limitation to the performance of a delta wing aircraft is vortex breakdown. Vortex breakdown occurs during high angle of attack flight, appearing initially at the rear of the wing and propagating toward the apex with increasing angle of attack. The jet-like velocity profile of the vortex, observed upstream of breakdown, is converted into a wake-like flow. This vortex breakdown is represented schematically in FIG.19, where it can be seen that the vortex 110 breaks down into a turbulent, wake-like flow 120. Downstream of the breakdown point, turbulence increases, axial velocities are substantially lower, and reverse flow is possible. Leibovich, S., "The Structure of Vortex Breakdown," Annual Review of Fluid Mechanics, Vol. 10, 1978, pp.45-88 (incorporated herein by reference). As velocity decreases, pressure increases resulting in a loss of lift. Furthermore, the location of vortex breakdown is known to fluctuate in the flow direction (typically at frequencies near 50 Hz). The fluctuations of the breakdown location, in addition to the turbulent

flow downstream of breakdown, result in significant buffeting and potential loss of stability.

[0050] Axial velocity profiles taken through the centerline of the vortex show the dramatic transformation that occurs as a result of breakdown. Novak, F. and Sarpkaya, T., "Turbulent Vortex Breakdown at High Reynolds Numbers," AIAA Paper 99-0135, 1999 (incorporated herein by reference). Well upstream of this location, the velocity profile is strongly jet-like with a peak velocity almost three times as great as the freestream velocity. Immediately downstream of breakdown the core flow stagnates. This rapid deceleration results in an increase in pressure and a decrease in lift. Interestingly, the deceleration of the core was found to begin upstream of the actual breakdown location. The consequences of this rapid deceleration are an increase in the pressure over the surface of the wing and the corresponding loss of lift. Campbell, J. F. "Augmentation of Vortex Lift by Spanwise Blowing," Journal of Aircraft, Vol. 14, No. 9, Sep. 1976, pp. 727-732 (incorporated herein by reference).

[0051] Although the exact cause of breakdown is still unknown, two major factors that seem to play a role are the swirl ratio and the presence of an adverse pressure gradient Leibovich; Dixon, C. J., "The Mechanism of Vortex Control by Spanwise Blowing and Wing Geometry," Lockheed-Georgia Co., Marietta GA, Engrg. Rpt. LG 78-ER-0187, June 1978 (incorporated herein by reference); Delery, JM, "Aspects of vortex breakdown," Progress in Aerospace Science, Vol. 30, 1994,. pp. 1-59 (incorporated herein by reference). Leibovich and others have shown that as the swirl ratio is increased, the location of breakdown moves upstream. In a delta wing, as the angle of attack increases, the level of swirl in the vortex increases. An adverse pressure gradient is also present over a delta wing due to the higher pressure at the trailing edge of the wing. As the angle of attack is increased, the pressure under the vortex becomes even lower, which creates a stronger adverse pressure gradient for the flow as it approaches ambient pressures at the trailing edge of the wing.

[0052] By delaying vortex breakdown (VBD), the wing's stall angle can be increased, thus increasing lift, and roll control can also be enhanced. By introducing powered blowing from control jets 130 near the delta wing apex, as shown in FIG.20,

the vortex core can be energized, thus delaying the VBD. As seen in FIG.19, the airstream 132 emitted by the control jet 130 air merges with the vortex 110, adding airflow to the vortex. This high velocity jet, which is emitted at the wing's apex where the vortex forms, provides additional momentum to the flow in the core. This provides the core with the energy needed to maintain stability against the adverse pressure gradients that are naturally present on the suction surface of the wing, pushing breakdown further downstream. This configuration also preserved the jet like profile of the vortex core over the entire length of the wing. The ability to maintain a high axial to tangential momentum ratio keeps the vortex coherent and prevents the development of conditions that lead to critical conditions that affect vortex instability and breakdown.

[0053] In our work, this blowing configuration provided significant improvement with good blowing momentum efficiency. It also proved to be rather sensitive to small changes in jet orientation, blowing momentum, and flight conditions. Experiments at off-design conditions showed that these slight changes had the ability to alter the trajectory of the jet such that it did not enter the core of the vortex as designed.

[0054] In order to delay VBD on a delta wing, pulsating injection by the control jets is preferable to continuous injection, with the preferred frequencies near 50 Hz corresponding to the characteristic time scale of VBD recovery. VBD control by pulsed injection has the advantage of requiring lower mass flow rate compared to continuous injection because the injection is not required all the time (partial duty cycle). These features make the PDE an ideal propulsion device for driving the control jets. The strong entrainment produced by the periodic exhaust of a PDE is advantageous in delaying the vortex breakdown due to the modification of the flow to a more favorable pressure gradient over the wing's surface.

[0055] Powered blowing at the trailing edge of the wing, downstream of the vortex breakdown, can also be used to delay vortex breakdown. As seen in FIG.21, this rear, downstream blowing can be achieved by diverting some of the PDE exhaust such that it exhausts in a rearward direction rather than being deflected around the trailing edge Coanda surface. The plenum chamber exhaust slots in the wing's trailing edge can be segmented to provide flexibility in providing a desired spanwise distribution of the

exhaust jets. FIG.21 shows one possible arrangement in which some of the jets are directed axially to provide thrust and VBD control, while the others wrap around the Coanda surface enhancing lift through supercirculation. Such re-direction of the PDE exhaust can be accomplished using the Coanda-effect canceling techniques discussed above and shown in FIGS.10 and 11. In this delta wing configuration, the streamwise component of the vortex flow provides propulsion while the low pressure induced on the wing surface by the high tangential and axial flow augments the aerodynamic forces. The geometry can adapt itself to speed and incidence by adjusting the momentum injected from the pulsed detonation. Pitching, rolling and yawing moments can be provided by differential power input by the PDE engines, frequency of detonation, fill fraction, or closing some of the distributed excitation along the leading edge.

[0056] We have observed that the optimum blowing arrangement used with this downstream nozzle works in a different manner than the apex control jets discussed above. Velocity measurements have shown that the downstream nozzle increased the axial velocity over the entire surface of the wing, not only in the vortex core but near the surface as well. In this configuration, the jet-like profile persisted only for a short distance downstream of the initial breakdown location and a wake-like profile developed, though with a higher axial core velocity than that observed with no control. The benefits of this increased velocity are lower surface pressures and an increased lift coefficient. Pressure was lower than the baseline value over the entire surface of the wing, particularly near the trailing edge where the wing has a larger span and more lift can be generated. The lift curve for the wing using the rear jet showed a substantial increase in lift at all angles of attack tested. An overall gain of 22% was achieved at the stall angle of 22°.

[0057] The rear nozzle generates improved flow by creating a low-pressure region at the rear of the wing. The flow velocity increase produced at the trailing edge by the control jet results in lower local static pressure. This low-pressure region creates a decreasing pressure pattern above the wing changing from high pressure near the apex to lower pressure at the aft region. This pressure variation is called "favorable pressure gradient" and it promotes greater axial flow in the vortex and delays breakdown. This mechanism is different than that observed in "trailing edge

blowing" which achieves the alleviation of adverse pressure gradient by increased entrainment.

[0058] Placement of the control jets on the trailing edge is not sensitive to injection angle or to the injection momentum because any jet orientation in the general downstream direction will provide flow acceleration and therefore reduced pressure. It is also not sensitive to the injection momentum because it does not have to penetrate into the vortex core as in the apex control jet case. Thus increased momentum will result in enhanced effect rather than penetration through the vortex without interaction as is the case in the apex jet.

[0059] Having described the invention with reference to exemplary embodiments, it is to be understood that the invention is defined by the claims and it is not intended that any limitations or elements describing the exemplary embodiment set forth herein are to be incorporated into the meanings of the claims unless such limitations or elements are explicitly listed in the claims. Likewise, it is to be understood that it is not necessary to meet any or all of the identified advantages or objects of the invention disclosed herein in order to fall within the scope of any claims, since the invention is defined by the claims and since inherent and/or unforeseen advantages of the present invention may exist even though they may not have been, explicitly discussed herein.

[0060] What is claimed is: