WO2017075718A1

WO2017075718A1 - Suppression of boundary layer separation in air-breathing engines

Info

Publication number: WO2017075718A1
Application number: PCT/CA2016/051290
Authority: WO
Inventors: Pradeep Dass; Craig JOHANSEN; E Jieh TEH
Original assignee: 1589549 Alberta Ltd.
Priority date: 2015-11-04
Filing date: 2016-11-04
Publication date: 2017-05-11
Also published as: CA3042947C; CA3042947A1

Abstract

An air breathing jet engine for a supersonic or hypersonic vehicle has an air passage having an air inlet, an air outlet, and a combustion chamber between the air inlet and the air outlet; one or more nozzles positioned in the air passage upstream of the combustion chamber; and a source of nanoparticles connected to the one or more nozzles to inject nanoparticles into the air passage.

Description

SUPPRESSION OF BOUNDARY LAYER SEPARATION IN AIR-BREATHING

ENGINES

TECHNICAL FIELD

[0001] This relates to injecting nanoparticles to enhance inlet performance in supersonic/hypersonic air-breathing engines, such as may be found in missiles, UAVs, aircraft, launch vehicles, etc., through the suppression of boundary layer separation caused by shock reflection. BACKGROUND

[0002] Flow separation induced by adverse pressure gradient induced by Shock Wave Boundary Layer Interaction (SWBLI) is a common phenomenon encountered in most supersonic/hypersonic capable flight vehicles. The phenomenon is usually undesirable and typically results in a degradation of vehicle performance. Numerous flow control devices have been developed to suppress this flow separation or reduce the size of the separation bubble generated, such as bleed hole, air injector (blowing), sub-boundary/micro vortex generator and more recently, arc plasma. All of these devices operate on the basic concept of increasing the momentum of the boundary layer and/or removing the lower momentum portion of the boundary layer. However, most of these flow control devices have their significant drawbacks that may indirectly degrade the performance of the flight vehicle. For example, bleed holes reduce the air flow rate into the engine, thus decreasing its operational efficiency, and need to be placed judiciously in the regions where separation will occur. Other methods may require additional design considerations in order to place and operate these devices, resulting in additional cost and increased weight of the vehicle.

SUMMARY

[0003] According to an aspect, there is described a flow separation suppression technique that uses seeded solid particles. When particles are injected into the flows of an air-breathing engine experiencing boundary layer separation, such as may occur in engines designed to operate at supersonic or hypersonic speeds, the separation bubble induced by the adverse pressure gradient from the shock wave is reduced. The rapid transfer of momentum between the particles and the gas at the separation location is critical for this invention. The magnitude of momentum transfer is controlled by the amount of particles deposited in to the flow while the rate of momentum transfer is controlled by the particle size. Sufficient rates of momentum transfer are only realized when the particles are at the nano-scale. To demonstrate the feasibility of the concept, a series of computational fluid dynamic (CFD) simulations were used to examine the effectiveness of the particles sizes and mass loadings on the suppression of the separation bubble may be examined. The CFD analysis employs high order numerical methods to capture shock wave structure and to eliminate artificial oscillations and Lagrangian Parcel Tracking algorithm to model and account for the solid particles. No turbulence model is used in the analyses as laminar flow assumption is invoked in all of the simulations, which represents higher altitude flight.

[0004] According to an aspect, there is provided an air breathing jet engine for a supersonic or hypersonic vehicle, the air breathing jet engine comprising an air passage having an air inlet, an air outlet, and a combustion chamber between the air inlet and the air outlet; one or more nozzles positioned in the air passage upstream of the combustion chamber; and a source of nanoparticles connected to the one or more nozzles to inject nanoparticles into the air passage.

[0005] According to an aspect, there is provided a method of reducing flow separation of air in the air passage during operation of an air breathing jet engine, the air breathing jet engine comprising an air passage having an air inlet, an air outlet, and a combustion chamber between the air inlet and the air outlet, the method comprising the step of: injecting nanoparticles into the air passage of the air breathing engine upstream of the combustion chamber, the nanoparticles being sized and injected at a rate sufficient to reduce flow separation of air in the air passage.

[0006] According to another aspect, the nanoparticles may be combustible.

[0007] According to another aspect, the nanoparticles may be sized and injected at a sufficient velocity to reduce flow separation of air in the air passage during operation.

[0008] In other aspects, the features described above may be combined together in any reasonable combination as will be recognized by those skilled in the art. [0009] While small sized (micron or nano) solid particles are frequently encountered in aerospace propulsion systems, this is predominantly in the combustion system and not with respect to suppressing flow separation. For example, solid rocket engines' exhausts contain solid particles as combustion products as well as soot particles formation in typical combustion system. As they are formed due to combustion processes, they usually do not possess any beneficial aspects in improving the performance of the propulsion system. More recently, there are emerging developments devoted to the use of solid particle additives in not only explosives, but in air-breathing supersonic/hypersonic propulsion systems. The recent induction of shock-induced combustion ramj et (scramj et) concept has also led to the inception of the use of powdered metals as fuel for such propulsion system.

BRIEF DESCRIPTION OF THE DRAWINGS

[0010] These and other features will become more apparent from the following description in which reference is made to the appended drawings, the drawings are for the purpose of illustration only and are not intended to be in any way limiting, wherein:

FIG. 1 is a schematic of an air-breathing engine designed to operate at hypersonic or supersonic speeds.

FIG. 2a is a graph depicting the velocity magnitude contours with a particle diameter of 16um.

FIG. 2b is a graph depicting the velocity magnitude contours with a particle diameter of 1.6um.

FIG. 2c is a graph depicting the velocity magnitude contours with a particle diameter of 160nm.

FIG. 2d is a legend for the graphs in FIG. 2a - 2c.

FIG. 3a is a graph depicting the boundary between the high and low velocity magnitude contours shown in FIG. 2a.

FIG. 3b is a graph depicting the boundary between the high and low velocity magnitude contours shown in FIG. 2b.

FIG. 3c is a graph depicting the boundary between the high and low velocity magnitude contours shown in FIG. 2c. DETAILED DESCRIPTION

[0011] Referring to FIG. 1, there is shown an example of a propulsion system that may be found on a scramjet (e.g. a ramjet in which combustion takes place in a stream of gas moving as supersonic speed) that incorporates the principles described herein. It will be understood that the principles herein may be applied to any type of air breathing jet engine 10. The depicted propulsion system includes a vehicle body 12 that defines a flow passage 14 that extends between an air intake 16 and a nozzle 18 at the air outlet 17 of flow passage 14 and has a combustion chamber 20. For simplicity, the fuel injection nozzles are not depicted. It will be understood that the principles may be used in other supersonic/hypersonic propulsion systems to improve performance.

[0012] As shown, multiple shock waves are generated on the vehicle forebody to compress the incoming air for combustion of the air-fuel mixture in the combustion chamber 20. This negates the need of a compressor, thus reducing the engine design requirement. The bow shock is shown at 30, while the shock-wave and boundary layer interaction occurs at 32.

[0013] In order to improve flow control of the shock waves and reduce the effect of flow separation, articles are injected into the flow passage 14 through particle injection nozzles 28 positioned in air passage 14 upstream of combustion chamber 20. Nozzles 28 are connected to at least one source of nanoparticles 29. Some benefit may be achieved by injecting particles at arbitrary points on the walls of the flow passage 14. However, beneficial result to the effectiveness of the flow control may be achieved by injecting particles a points upstream of the SWBLI region, or prior to where the flow separation will occur, such as the location indicated by reference number 22. As the injection of the particle suspension will include the introduction of additional air into the main flow, this will be similar to other flow control techniques, i.e. blowing, which will increase the momentum of the boundary layer. Such embodiments may lead to unnecessary additional use of amount of particles to suppress the separation. Furthermore, multiple injection points will further increase the complexity of the engine design. [0014] The location at which the particles may be injected may be selected in order to achieve the maximum benefit or may be variable. This may be done in order to account for the different speeds of the flight vehicle, which may vary the locations of the shock impingement. For example, the particles may be injected at location 24, near the true intake of the combustion chamber 20. When the particles are introduced into the flow in such a manner, the transverse jet into supersonic crossflow (TJISC) will create counter-rotating vortices. These vortices will aid in increasing the distribution of the particles prior to reaching the shock train 26 where complications arising from SWBLI are severely compounded. [0015] In some circumstances, it may be beneficial to provide a mass flow rate of particles that is at least about 10% of the inlet air intake.

[0016] Although seeding solid particles could present a serious weight penalty to the flight vehicle, these particles may optionally function as part of the fuel requirement for the vehicle, by providing nanoparticles that are combustible. The particles may be selected with high energy densities, such as certain metallic particles, which may have higher energy densities than pure hydrocarbons alone that can be released through combustion. Furthermore, solid particles can be more easily stored and have small storage volume requirements. [0017] The transfer of momentum between the nano-particles and the air at the separation location is the mechanism that leads to the suppression of flow separation. At the location of the shock wave 30, the gas experiences a sudden decrease in velocity that occurs over a small distance (i.e. thickness of the shock wave). Since the injected nano-particles have inertia they have a higher velocity than the gas in the post-shock region 32. The mismatch in gas and particle velocities in the post-shock region leads to particle drag. Through aerodynamic drag, the momentum of the nano-particles is transferred to the gas. When this occurs at the wall, the boundary layer becomes energized and is able to resist the adverse pressure gradient that normally leads to flow separation. As will be understood by those skilled in the art, the nanoparticles may be a variety of sizes and may be injected at varying velocities. It will be understood that the nanoparticles may be sized such that they can be injected at sufficient velocity to reduce flow separation of air in air passage 14 during operation. [0018] FIG. 2a - 2c show some of the CFD results of the effect of nano-particles on the SWBLI. In these simulations, the freestream flow is at a Mach number of 2.15, a stagnation pressure of 10.9 kPa, and a stagnation temperature of 300K. An oblique shock with a half wave angle of 33.18° impinges on to a laminar boundary layer. In the figure, contours of gas velocity magnitude indicate where there is separation (i.e. low velocity indicates separation). The particle mass loading, or the ratio of the total particle mass to the air mass, was fixed at 0.1. The results indicate that the effectiveness of the particles to suppress separation increases as the particle diameter decreases. A substantial suppression of the separation only occurs when the particles are at the nano-scale. FIG. 3a - 3c depict the boundary 34 between the high velocity conditions 36 and the low velocity conditions 38 of FIG. 2a - 2c to clarify the relationship.

[0019] In this patent document, the word "comprising" is used in its non-limiting sense to mean that items following the word are included, but items not specifically mentioned are not excluded. A reference to an element by the indefinite article "a" does not exclude the possibility that more than one of the elements is present, unless the context clearly requires that there be one and only one of the elements.

[0020] The scope of the following claims should not be limited by the preferred embodiments set forth in the examples above and in the drawings, but should be given the broadest interpretation consistent with the description as a whole.

Claims

What is Claimed is:

1. An air breathing jet engine for a supersonic or hypersonic vehicle, the air breathing jet engine comprising:

an air passage having an air inlet, an air outlet, and a combustion chamber between the air inlet and the air outlet;

one or more nozzles positioned in the air passage upstream of the combustion chamber; and

a source of nanoparticles connected to the one or more nozzles to inject nanoparticles into the air passage.

2. The air breathing jet engine of claim 1, wherein the nanoparticles are combustible.

3. The air breathing jet engine of claim 1 or 2, wherein the nanoparticles are sized and injected at a sufficient velocity to reduce flow separation of air in the air passage during operation.

4. A method of reducing flow separation of air in the air passage during operation of an air breathing jet engine, the air breathing jet engine comprising an air passage having an air inlet, an air outlet, and a combustion chamber between the air inlet and the air outlet, the method comprising the step of:

injecting nanoparticles into the air passage of the air breathing engine upstream of the combustion chamber, the nanoparticles being sized and injected at a rate sufficient to reduce flow separation of air in the air passage.

5. The method of claim 4, wherein the nanoparticles are combustible.