US3225589A

US3225589A - Apparatus for testing the principles of detonation combustion

Info

Publication number: US3225589A
Application number: US101800A
Authority: US
Inventors: Jr Selden B Spangler; James L Munier
Original assignee: Garrett Corp
Current assignee: Garrett Corp
Priority date: 1961-04-10
Filing date: 1961-04-10
Publication date: 1965-12-28
Anticipated expiration: 1982-12-28

Description

Dec. 28, 1965 s N LE JR" ET AL 3,225,589

APPARATUS FOR TESTING THE PRINCIPLES OF DETONATION COMBUSTION Filed April 10, 1.961

INVENTORS SELDE/V B. SPANGLER JR. BY JAMES L. I??? ATTORNEY United States Patent Ofiice 3,225,589 Patented Dec. 28, 1965 APPARATUS FOR TESTING THE PRINCIPLES OF DETONATIGN CQMBUSTION Selden B. Spangler, Jr., and James L. Munier, Scottsdale,

Ariz., assignors to The Garrett Corporation, Los Angeles, Calif., a corporation of California Filed Apr. 10, 1961, Ser. No. 101,890

9 Claims. (Cl. 73-35) This invention relates to test apparatus, and more particularly to an apparatus for simulating the air flow and high altitude conditions of hypersonic flight for the purpose of testing the principles of detonation combustion in hypersonic aircraft.

Recently, considerable work has been done on the development of a variable geometry detonation combustion engine for use in hypersonic aircraft. Such an engine and its theory and method of operation are disclosed in the copending application of Hunter and Norman, Serial No. 88,149, filed February 9, 1961. combustion process as described by Hunter and Norman, the flame front or detonation is established and maintained in a variable geometry aerothermodynamic duct or engine by a shock wave which has the characteristic that a static temperature rise occurs in a very short distance across the wave. The detonation is stabilized in the engine by causing the shock wave to be maintained in a fixed position relative to the confining structure, and subsequent expansion of the gaseous detonation products is used to develop a continuous thrust for the hypersonic aircraft.

In order to study the various aspects of the detonation combustion process and the many variables encountered in supersonic and hypersonic flight, it is desirable to have a test tunnel in which the anticipated operating conditions may be simulated. While work has been done on the detonation combustion process in test tunnels, any such tunnels that have been available in the past have had certain deficiencies. For example, prior detonation combustion tunnels have not been designed to reproduce flight conditions for which the detonation combustion process has been shown to have most promise. In particular, the combustion air total temperature conditions have not been simulated; nor have fuel injection conditions been simulated. In addition, oxygen from the combustion air has been used in conjunction with a fuel for preheating the combustion air stream, with the result that there is a deficiency of oxygen in the detonation combustion zone. The present invention is based, at least in part, on the discovery that when hydrogen is used as the fuel in a test tunnel for the detonation combustion process, most prior difliculties may be obviated by preheating the combustion air stream with a flame derived from controlled outside sources of hydrogen and oxygen. In addition, preignition of the detonation fuel may be prevented and fuel injection more adequately studied by adding the fuel at a point or zone in the supersonic portion of the air stream at which the static temperature is below the ignition temperature of hydrogen in air. It has also been found that the overall size of the entire apparatus may be minimized by utilizing bleed air from a gas turbine as the source of high pressure combustion air.

It is an object of this invention to provide an improved apparatus for simulating the flow and altitude conditions of hypersonic flight for testing detonation combustion.

Another object of this invention is to provide a test apparatus for simulating hypersonic flight fluid flow for detonation combustion in which a stream of combustion air is supplied from a high speed gas turbine.

Another object of this invention is to provide a test apparatus for simulating hypersonic flight fluid flow for detonation combustion in which the combustion air In the detonation d stream is preheated with a gas flame derived from controlled outside gas sources so as to maintain a predetermined oxygen supply in the air stream.

A further object of this invention is to provide a test apparatus for simulating hypersonic flight fluid flow for detonation combustion in which fuel is added in the supersonic portion of the air stream where the static temperature is below the ignition temperature of hydrogen in air.

It is another object of this invention to provide a test apparatus for simulating hypersonic flight fluid flow for detonation combustion in which the fuel injector is streamlined so as to reduce any disturbance in the air stream to a minimum.

A still further object of this invention is to provide a test apparatus for simulating hypersonic flight fluid flow for detonation combustion in which the geometry and location of the preheater and fuel injector in the apparatus are such as to provide the proper temperature and fuel distribution over an appreciable portion of the detonation combustion zone, which portion does not extend to the walls of the apparatus so that there will be no deleterious effects on the walls due to high temperature.

The above and other features and objects of the invention will be apparent from the following description and the accompanying drawing, in which:

FIG. 1 is a schematic side elevational view of a supersonic testing apparatus embodying the features of this invention;

FIG. 2 is a longitudinal vertical sectional view of portions of the preheater and tunnel sections of the apparatus;

FIG. 3 is a horizontal sectional view of the preheater burner, taken along the line 33 of FIG. 2; and

FIG. 4 is a horizontal sectional view of the fuel injector, taken along the line 4-4 of FIG. 2.

Referring now to the drawing, and particularly FIG. 1, it will be observed that the test apparatus in which the features of this invention have been incorporated comprises a source of high pressure air 10 from which the air is directed through a stilling chamber 11 where its velocity is reduced, turbulence eliminated, and the air stream is rendered substantially homogeneous as it leaves said chamber. The stilled air stream then passes through a preheater 12 to an inlet 13 of a test tunnel or elongated aerothermodynamic duct 14 having an exhaust 15. In the present instance, the duct 14 is shown as rectangular in cross section, though obviously it could have any desired cross section, such as polygonal, elliptical or circular. The test apparatus is designed particularly for simulating the operating conditions of a detonation combustion engine in hypersonic flight. While any convenient source of high pressure air may be used, it has been found to be advantageous to utilize the high pressure bleed air stream produced by a gas turbine 16, since this reduces the size and weight of the overall apparatus. As is well known, such turbines usually include a compressor section 17 adjoining a combustor section 18 which leads to the turbine 16, the burned gases being discharged through an exhaust 29. In this instance, a conduit 21 connected to the compressor section 17 leads the high pressure bleed air stream to the stilling chamber 11 wherein its velocity is reduced and flow inhomogeneities are minimized. The substantially homogeneous flow is then ducted to the preheater 12. In the event that electric power may be needed in the test apparatus, such as for operating thermocouples and other test instruments, the gas turbine 16 could obviously be used to drive a suitable electric generator (not shown).

Prior to the present invention, there has been some question, in the use of test tunnels of this general character, as to the variation of oxygen concentration in the detonation combustion zone and the concomitant effect of such variation on the characteristics of the combustion process because of the use of the oxygen in the combustion air in combination with a fuel to preheat the air. According to this invention, the air stream is preheated with a flame produced by independent and separately controlled sources of hydrogen and oxygen. As shown in FIGS. 2 and 3, the preheater 12 includes a housing 22 to which the air conduit 21 is connected, and air from the compressor 17 thus passes through the preheater 12 into the inlet 13 of the test tunnel. A preheater burner 23 is suitably mounted in the housing and provided with a starter section 24 which includes a glow-plug 25 connected to a convenient source of power 26. The burner 23, in the form shown, comprises a streamlined or wedgeshaped upstream end 27 having a relatively sharp leading edge 28 on the upstream side thereof and a vertically disposed groove 30 on the downstream side or edge thereof. A central burner opening 31 is positioned in the groove 3-9 in a vertical position corresponding approximately to the center of the air stream, and this opening is connected by a passage 32 and line or conduit 33 to a pressure source or tank 34 of hydrogen gas. A valve 35 downstream of a pressure regulator 36 in the line 33 may be used to control the flow of fuel to the burner opening 31. Oxygen required for combustion of the fuel is directed to the burner area through a plurality of openings 37 which may be arranged in a circle around the opening 31. The openings 37 are connected through passages 38 to a conduit or pipe line 39 which leads to a source or tank 40 of oxygen under the control of a flow control valve 41 and a pressure regulator 42. It will be understood that, if desired, the gases could be reversed in the burner so that oxygen would be supplied to the opening 31 and hydrogen to the surrounding openings 37.

The starter section 24 has connected to it a conduit 43 which leads to a mixer 44. Hydrogen and oxygen are supplied to the mixer 44 by conduits 45 and 46, respectively, which emanate from the tanks 34 and 40. Pressure regulator 47 and flow control valve 48 regulate the hydrogen flow to the mixer 44; and similarly pressure regulator 5t) and flow control valve 51 regulate the flow of oxygen from tank 40 to said mixer. When it is desired to start the preheater burner 23, hydrogen and oxygen are first supplied to the mixer 44 and the mixed gases flow through conduit 43 to the igniter section 24. The glowplug 25 is supplied with electric current and causes the gas mixture flowing into groove 30 from conduit 43 to be ignited. The flame moves along the groove to the

openings

31 and 37 which are then supplied with hydrogen and oxygen through

conduits

33 and 39 and these gases mix and are ignited. The glow-plug 25 and igniter gas flow may then be turned off.

By having a separate supply of oxygen for the preheater, as described, proper amounts of oxygen may be supplied to the burner 23 so that the products of combustion, which are water vapor and excess oxygen, will have the same oxygen-to-inerts (water vapor) ratio as that or" oxygen-to-inerts (nitrogen plus traces of other noncombustibles) in normal unburned air. In this manner, the preheat process will have the least effect on the oxygen concentration at the detonation combustion zone. In addition, the oxygen is most thoroughly mixed in the air-water vapor stream because of its direct use in the preheat combustion process. The design and operation of the preheater are such that a large proportion of the air stream entering the test tunnel is heated to predetermined temperatures as high as 3500 F.

Detonation combustion, which will produce optimum thrust conditions for hypersonic flight in the neighborhood of Mach 6.5, requires supersonic air flow in the detonation combustion zone of about Mach 3 and a temperature of 3400 F. To increase the velocity of the heated air stream as it flows from the preheater into the tunnel inlet 13, a nozzle 52 having a restricted throat 53 is provided in the tunnel 14 adjacent said inlet 13. The internal geometry of the tunnel also includes special wedges 54 on the top and bottom inside surfaces toward the downstream end of the tunnel and these act to hold a normal shock wave 55 in the desired position in the tunnel, as explained in the Hunter and Norman application referred to above. It is across this normal shock wave that detonation combustion takes place when fuel (in this instance, hydrogen) is supplied to the tunnel.

A special streamlined fuel injector 56 is provided to supply hydrogen fuel to the tunnel and inject such fuel into the supersonic air stream. This injector, which in the form shown is diamond-shaped, comprises an upstream Wedge-shaped portion 57 having a sharp leading edge 58 and an adjoining downstream wedge-shaped section 60. These two wedge-shaped sections having their bases contiguous to one another provide the desired streamlined shape for the injector 56, which is shown as vertically disposed substantially midway between the sides of the duct or tunnel in a position downstream of the nozzle 52 in a uniform flow field but upstream of the shock wave 45. It will be understood that the injector could be horizontally disposed midway between the top and bottom of the tunnel, if desired. Fuel may be injected into the passing supersonic air stream through an opening or nozzle 61 provided in the downstream edge 62 of the injector. Nozzle 61 is connected by a passage 63 and a conduit 64 to the hydrogen tank 34 and is under the control of a flow control valve 65 and a pressure regulator 66.

To cool the fuel injector, and thus aid in preventing preignition of the fuel, a cooling fluid inlet passage 67 is formed in the upstream wedge-shaped portion 57 adjacent the leading edge 58. This passage is connected by a conduit 68 to a source 70 (FIG. 1) of coolant. Passage 67 is connected with a centrally disposed return passage 71, which in turn is connected to a conduit 72 leading to a sump (not shown) to collect the used cooling fluid. A pump 73 and valve 74 control the flow of cooling fluid, which in this instance is water, from the tank or source 70 to the fuel

injector cooling passages

67 and 71 and also to a conduit 75. This latter conduit leads to a cooling header 76 disposed over the length of the tunnel 13 and provided with a plurality of sprinkler nozzles 77. In this way the entire tunnel may be water-cooled.

In the event that it is desired to cool the tunnel nozzle 52, cooling passages 78 may be provided around the throat 53, as shown in FIG. 2. A pipe or conduit 80 is then connected to such cooling passages and to the source of cooling fluid. In this instance, the conduit 80 is shown in FIG. 1 as being connected to the header 76.

Since one purpose of the apparatus is to study the detonation wave, observation windows 81 may be provided on each side of the tunnel 14 at a position adjacent the wedges 54.

By means of the test apparatus described above, it is possible to study the many variables that are involved in the detonation combustion process, and also to study such variables and conditions as they may be encountered in a variable geometry engine under changing flight conditions. One of the principal and important problems in such an engine is the method of injecting and controlling the fuel flow so as to effect minimum specific fuel consumption. Consequently, the regulation and control of the oxygen in the air stream to simulate air conditions at flight altitude is important and according to this invention is possible with the specially controlled preheater.

Various changes may be made in the construction and certain features may be employed without others Without departing from the invention or sacrificing any of its advantages.

We claim:

.1. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying air at high pressure to said duct; means for preheating such air while at the same time regulating the amount of oxygen therein; means for establishing and maintaining a shockwave in said duct; and means for injecting fuel into the airstream in said duct so that detonation may take place across such shockwave.

2. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; a gas turbine for supplying an air stream at high pressure to said duct; means between said turbine and duct for preheating said air stream while at the same time controlling the amount of oxygen in said air stream; means for establishing and maintaining a shockwave in said duct; and means for injecting fuel into the airstream in said duct so that detonation may take place across such shockwave.

3. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying air at high pressure to said duct; means for preheating such air while at the same time regulating the amount of oxygen therein; means for creating a normal shock wave in said duct; and means upstream from said shock Wave for injecting controlled amounts of fuel into the preheated supersonic air stream.

4. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying an air stream at high pressure to said duct; means for preheating said air stream while at the same time regulating the amount of oxygen therein; means for accelerating said air stream to supersonic velocities; and streamlined means in said duct for injecting controlled amounts of fuel into the preheated air stream after acceleration of said air stream to supersonic velocities.

5. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying an air stream at high pressure to said duct; means for preheating said air stream While at the same time regulating the amount of oxygen therein; means for accelerating said air stream to supersonic velocities; streamlined means in said duct for injecting controlled amounts of fuel into the preheated air stream after acceleration thereof to supersonic velocities; and means for circulating cooling fluid through such fuel injecting means.

6. A test apparatus for simulating supersonic air flow and high altitude conditions for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying an air stream at high pressure to said duct; means for stilling said air stream; means for preheating said air stream while at the same time regulating the amount of oxygen therein; means for accelerating said air stream to supersonic velocities; streamlined means in said duct for injecting controlled amounts of fuel into the preheated air stream after acceleration thereof to supersonic velocities; and means for cooling the outside of said duct and the inside of such fuel injecting means.

7. A test apparatus for simulating supersonic air flow and high altitude conditions for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying an air stream at high pressure to said duct; means for stilling said air stream; means for preheating said air stream While at the same time regulating the amount of oxygen therein; means for accelerating said air stream to supersonic velocities; streamlined means in said duct for injecting controlled amounts of fuel into the preheated air stream after acceleration thereof to supersonic velocities; and means for cooling the air stream accelerating means and the inside of such fuel injection means.

8. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying air at high pressure to said duct; means for preheating such air while at the same time regulating the amount of oxygen therein; means for creating a normal shock Wave in said duct; and means upstream from said shock Wave for injecting controlled amounts of fuel into the preheated supersonic air stream, the geometry and location of the preheating means and the fuel injecting means being such as to provide proper temperature and fuel distribution over an appreciable portion of said normal shock wave Where detonation combustion takes place, such portion being of limited extent so as to prevent harm to the walls of the duct due to high temperature.

9. A test apparatus for simulating the air flow conditions in hypersonic flight for testing the principles of detonation combustion, comprising: an elongated aerothermodynamic duct; means for supplying air at high pressure to said duct; means for preheating such air, said preheating means including means for supplying fuel thereto and supplying regulated amounts of oxygen to said preheating means and said duct; means for creating a normal shockwave in said duct; and means upstream from said shockwave for injecting controlled amounts of fuel into the preheated supersonic air stream.

References Cited by the Examiner UNITED STATES PATENTS 2,615,331 10/1952 Lundgren 73-116 2,763,155 9/1956 Beams et al. 73116 2,885,890 5/1959 Liccini et al. 73--147 3,005,338 10/1961 Libby et al 73147 OTHER REFERENCES Cushman, R. H; Aviation Week, January 6, 1958, pages 57, 58, 59, 63.

Dugger, G. L.: ARS Journal, November 1959, pages 819-827.

LOUIS R. PRINCE, Primary Examiner.

DAVID SCHONBERG, Examiner.

Claims

1. A TEST APPARATUS FOR SIMULATING THE AIR FLOW CONDITIONS IN HYPERSONIC FLIGHT FOR TESTING THE PRINCIPLES OF DETEONATION COMBUSTION, COMPRISING: AN ELONGATED AEROTHERMODYNAMIC DUCT; MEANS FOR SUPPLYING AIR AT HIGH PRESSURE TO SAID DUCT; MEANS FOR PREHEATING SUCH AIR WHILE AT THE SME TIME REGULATING THE AMOUNT OF OXYGEN THEREIN; MEANS FOR ESTABLISHING AND MAINTAINING A SHOCKWAVE IN SAID DUCT; AND MEANS FOR INJECTING FUEL INTO THE AIRSTREAM IN SAID DUCT SO THAT DETONATION MAY TAKE PLACE ACROSS SUCH SHOCKWAVE.