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Publication numberUS3646761 A
Publication typeGrant
Publication dateMar 7, 1972
Filing dateMar 7, 1961
Priority dateMar 7, 1961
Publication numberUS 3646761 A, US 3646761A, US-A-3646761, US3646761 A, US3646761A
InventorsHunter Skillman C, Norman Leslie W, Perrone George L
Original AssigneeGarrett Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Method and apparatus for starting detonation combustion engines
US 3646761 A
Abstract
1. In a method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet and an adjustable throat adjacent the combustion zone and is adapted to receive a hypersonic stream of air during normal operation, the steps comprising reducing the amount of air supplied by the inlet to the throat while simultaneously opening the throat until at least sonic velocity of the fluid stream is attained at the combustion zone, then gradually increasing the amount of air supplied by the inlet to the throat, feeding fuel into the fluid stream, and restricting the throat so as to adjust the flow to obtain the desired velocity, temperature and pressure for detonation in said combustion zone.
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United States Patent Norman et al.

[ Mar. 7, 1972 [72] Inventors: Leslie W. Norman, Scottsdale; Skillman C. Hunter; George L. Perrone, both of Phoenix, all of Ariz.

[73] Assignee: The Garrett Corporation, Los Angeles,

Calif.

[22] Filed: Mar.7,1961

[2]] Appl. No.: 94,022

OTHER PUBLICATIONS Cushman Scientists Study Mach 7 Ramjet Theory," Aviation Week, Vol. 68, No. 1, Jan. 6, 1958- pages 57 59 and 63 Primary Examiner-Robert F. Stahl Attorney-Herschel C. Omohundro and John H. G. Wallace EXEMPLARY CLAIM 1. In a method of starting a detonation combustion engine having an aerothennodynamic duct which includes an adjustable inlet and an adjustable throat adjacent the combustion zone and is adapted to receive a hypersonic stream of air during normal operation, the steps comprising reducing the amount of air supplied by the inlet to the throat while simultaneously opening the throat until at least sonic velocity of the fluid stream is attained at the combustion zone, then gradually increasing the amount of air supplied by the inlet to the throat, feeding fuel into the fluid stream, and restricting the throat so as to adjust the flow to obtain the desired velocity, temperature and pressure for detonation in said combustion zone.

8 Claims, 7 Drawing Figures 0000 0 0 0. .0. 0000 0.000 0 00 0 m mw 0000 60 0 0000 0000 0 00 0 00 0 mmw 0000 00 0 000. 0000 0000 0 0 02 t R R .0025 0000 00 0 0.00. 00000 0 b0. 0 0 000 00 0 00: 0000; 0 00. 0 mmw M 0000 2 0 0. .0 0000 0 00. 0 MM W 0 000 00 0 05 003 0 00 0. 5 0% x 02.0 00 0 000 h 0000 0 0.2 0. J 0 000 00 0 000. 0000 .000 0 0 \A 9 6%: E: E0 E0 020: 2.0% OU 0 0 W 0. 0 0% 00 000 5 2%: ML.

PATENTEUMAR 7 I972 A TTORNE'Y METHOD AND APPARATUS FOR STARTING DETONATION COMBUSTION ENGINES This inventionrelates to airbreathing engines for aircraft, and aims to provide a method of and means for controlling the intake and airflow during the starting of a detonation combustion engine used in hypersonic aircraft.

In the copending application of Hunter and Norman, Ser. No. 88,149 filed Feb. 9, 1961, a detonation combustion engine is disclosed which is designed specifically for :use in hypersonic aircraft. Accordingto the Hunter and Norman application, the flame front or detonation is established and maintained by positioning and maintaining a stable aerodynamic shock system in a variable geometry aerothermodynamic ductor engine such that the-heat release occurs across a shock wave, rather than being maintained by a flame holder as is the case in conventional internal combustion processes such as occur in ramjet and turbojet engines. Steady state detonation is established and maintained across a standing shock wave of fixed position relative-to a confining structure therefor, and subsequent expansion of the gaseous detonation product is used to develop a continuous thrust.

Starting the detonation combustion process in a detonation combustion engine having an aerothermodynamic duct presents many of the same problems that are encountered in wind tunnels and supersonic inlets. Immediately upon opening the duct or engine inlet to a hypersonic (Mach number above 5.0) free stream, the conditions inside the duct, being subsonic, are such that the air captured by the inlet cannot be completely passed through it. As a result, a detached normal shock wave forms outside the inlet lip in such a position as to allow sutficient air to spill out the sides that the remainder passes through the duct at a subsonic rate which is thus inadequate for the production of the desired conditions for detonation combustion.

The present invention is based on the discovery that in a variable geometry detonation combustion engine, the abovementioned starting difficulties may be-completely eliminated by diverting a sufficient portion of the inlet air, while at the same time opening the throat immediately adjacent the detonation combustion zone, that the captured air may proceed through the engine to the detonation combustion zone at the desired supersonic velocity. When this result is attained, the amount of air allowed into the engine atthe inlet may be increased to the full stream striking the inlet, and subsequently the throat immediately adjacent the zone of detonation may be restricted so as to adjust the levels of pressure, temperature and velocity to the operating condition most favorable for detonation with minimum specific fuel consumption for the thrust required.

It is therefore an object of this invention to provide an efficient method of and means for starting a detonation combustion engine designed for use in hypersonic aircraft.

Another object of this invention is to provide an efficient method of starting a variable geometry detonation combustion engine in which the amount of air entering the engine and the rate of flow through the engine are adjusted to effect ideal conditions for detonation.

A further object of this invention is to provide a method of starting a detonation combustion engine having an adjustable inlet and an adjustable throat adjacent the combustion zone, which method comprises reducing the amount of air entering the inlet while simultaneously opening the throat until a predetermined velocity of the fluid stream is attained at the combustion zone, then gradually increasing the amount of air entering the inlet and reducing the throat area to obtain the desired velocity, temperature and pressure for detonation in said combustion zone when a suitable fuel is injected into the fluid stream.

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

FIG. I is a schematic side elevational view of a hypersonic aircraft having a detonation combustion engine embodying the principles of the invention;

FIG. 2 is a schematic plan view of the detonation combustion engine shown in FIG. 1;

FIG. 3 is a table of the values of pressure, temperature,

-Mach number, and velocity prevailing in the various flow repersonic propagation velocity and a large pressure and temperatureincrease across the wave. A detonation differs from a subsonic flame in that such a flame moves with subsonic velocity andits microscopic propagation mechanisms are fundamentally different. The supersonic combustion wave produces strong detonation which is stable, steady, reproducible, and obtainable over a wide fuel-air-ratio as long as the Mach number of the approach flow is'greater than the Chapman-Jouguet Mach number for that fuel-air ratio. A Chapman-Jouguet" detonation is characterized by the fact that the flow immediately behind the wave is sonic, i.e., a Mach number of one, and such a detonation represents the minimum supersonic propagation Mach number for a given fuel-air ratio.

Experiment has shown that detonation combustion takes place under the following conditions:

I. A stream of gases is moving supersonically relative to the containment vessel or aerothermodynamic duct;

2. The total temperature of the supersonic gas stream must exceed the detonation temperature of the gas stream;

3. The gas stream consists of a mixture of some oxidizer, such as air, and unburned fuel; such as hydrogen; and

4. A shock system of such strength is generated at some point in the aerothermodynamic duct that the static temperature across the shock system is high enough to cause detonation.

This type of detonation can occur across a normal shock wave whence the Mach number downstream of the shock wave will be subsonic. The same phenomena may occur across an oblique shock wave as long as the static temperature downstream of the shock wave is sufficiently high to cause detonation.

Referring now to FIGS. 1 and 2, a hypersonic aircraft A, shown schematically, is powered by a variable geometry detonation combustion engine E to' produce hypersonic speeds in the range of 6 to 7 Mach number at an altitude of 125,000 feet. The design of the airframe and the inlet of the engine are such as to produce a series of shock wave 8,, S S S S and S which will slow the captured flow from a free stream Mach number of 6.5 to a Mach number of 2.5 at the inlet to the detonation zone. The manner of producing the shock waves to accomplish this reduction in Mach number, with an attendant increase in static pressure and static temperature, is fully described in the copending Hunter and Norman application referred to above. Briefly, the design of the airframe is such that first and second critical shock waves S and S: are produced upstream of the inlet to the engine E. These waves define regions 12 and 13 wherein the pressure, temperature, velocity, and Mach numbers have the values shown in the table of FIG. 3. Region 11 is the free stream adjacent the aircraft. The geometry of the engine is such that shock waves 8;, and S are formed and create regions 14, I5 and 16 where the physical conditions are as tabulated in FIG. 3. Next, critical shock wave S (FIG. 1) is created by the hinged leading edge of an adjustable ramp or door 20 which is arranged in the combustion zone immediately upstream from the normal shock wave 5,. This last wave S, is the one across which detonation combustion takes place. Wave S, creates zone 17, immediately prior to combustion, and the physical conditions in this zone are at least partially controlled by the angle of the adjustable ramp, as will be described more fully hereinafter.

The table of FIG. 3 shows that a stepwise increase in static temperature accompanies each successive shock wave in the aircraft and engine duct system. To take advantage of this increase, a fuel, such as hydrogen, having a detonation temperature below the static temperature occurring in the normal shock wave S, is recommended. The FIG. 3 table also indicates that the statictemperature occurring in the flow region 17 downstream of oblique shock wave 8,, may be higher than the temperature required for initiation of combustion in the air-hydrogen mixture. The time required for this mixture to traverse region 17 is so short, however, in comparison to the time needed to achieve combustion, that the mixture may remain chemically unafl'ected until it reaches the normal shock wave 8,. Here (region 18) the static temperature rises with consequent detonation of the mixture and a further increase in static temperature to 4,950R., as shown in FIG. 3 Finally, region 19, at the rear of the exhaust, indicates the con- .ditions of thrust.

The foregoing explanation of the shock wave system pertains primarily to the operation of the engine during cruising. As explained in the Hunter and Norman application above referred to, the aircraft may be accelerated to a hypersonic cruising speed of about Mach number 6.5, for example, by a jettisonable booster rocket (not shown). During this initial period of acceleration, the inlet 21 of engine E is closed by a removable door 22 which also may be jettisoned when it is desired to start the engine in accordance with the principles of this invention.

Assuming now that the aircraft has been accelerated to the desired cruising speed and the removable door has been jettisoned, at this moment, the air inside the aerothermodynamic duct of the engine will be substantially stationary or subsonic. When the full force of the airstream then strikes the inlet 21, the air captured by the inlet cannot be completely passed through it because the air in the duct is subsonic. As a result, a normal shock wave 5, forms outside the inlet lip, as indicated in FIG. 4, in such a position as to allow large amounts of air to spill out the sides of the inlet 21. It has been found that the amount spilled is such that the remainder then passes through the engine duct at a subsonic rate which is thus inadequate for the production of the desired conditions for detonation combustion.

To eliminate this difficulty and permit the development of the desired supersonic velocities in the engine duct, as described above and indicated in the table of FIG. 3, the procedure now to be described and forming the subject matter of this invention is followed. In FIGS. 1 and 4-7, it will be noted that a special movable door'23 is provided in the lower wall or floor of engine E near or immediately downstream of the inlet 21. This door 23 is normally flush with the bottom of the duct and functions in cooperation with the adjustable ramp 20 during the starting operation.

As shown in FIG. 4, when the removable cover 22 is jettisoned, the door 23 is closed and the ramp 20 is moved to the open position shown by any conventional means, such as a pneumatic piston 24 which is arranged to receive fluid pressure from a suitable source 25. When the ramp 20 is in this open position, the throat of the engine duct immediately adjacent the combustion zone is wide open and substantially unrestricted. In spike of this, however, the normal shock wave 8,, will stay outside the inlet if the doors 23 and 20 remain in the FIG. 4 position. According to this invention, part of the captured air is bypassed by opening the inlet door 23 to the FIG. position, and this permits the normal shock wave 8,, to move downstream past the inlet lip as indicated. Inlet door 23 may be moved to open position in any desired manner, such as by a second pneumatic piston 26 also supplied with fluid under pressure from source 25.

Having thus moved inside the inlet, the normal shock wave S, proceeds on downstream, with the establishment of suitable operating conditions, until it reaches the position shown in FIG. 1 where it is retained by steep angled wedges 27 (FIG. 2). This is the position in which detonation combustion takes place following the injection of predetermined amounts of fuel, such as hydrogen. The fuel may be fed from a supply tank 28 to a plurality of injectors 30 which are positioned in the engine duct upstream of the combustion zone, such as in region 14. Prior to the injection of fuel, however, the amount of bypassed air is decreased by closing the door 23 gradually (FIG. 6); and substantially at the same time adjusting the angle of the combustor ramp upwardly to some position between those shown in FIGS. 6 and 7, depending upon the specific conditions desired. The adjustability of the angle of the ramp 20 makes it possible to regulate the velocity, temperature and pressure to those suitable for detonation, so that when fuel is then added, as described above, such detonation will take place at the FIG. 1 position of wave 8,. When the ramp 20 is in its fully open position, rendering the throat adjacent the combustion zone substantially unrestricted, the contraction ratio of the engine duct is decreased, resulting in the required reduction of the amount of bypassed air to effect the desired downstream movement of the normal shock wave S3.

Various changes may be made in the method and means described herein, and certain features may be employed without others, without departing from this invention or sacrificing any of its advantages.

We claim:

1. In a method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet and an adjustable throat adjacent the combustion zone and is adapted to receive a hypersonic stream of air during normal operation, the steps comprising reducing the amount of air supplied by the inlet to the throat while simultaneously opening the throat until at least sonic velocity of the fluid stream is attained at the combustion zone, then gradually increasing the amount of air supplied by the inlet to the throat, feeding fuel into the fluid stream, and restricting the throat so as to adjust the flow to obtain the desired velocity, temperature and pressure for detonation in said combustion zone.

2. A method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet and an adjustable throat adjacent the combustion zone and is adapted toreceive a hypersonic stream of air during normal operation, which method comprises reducing the amount of air supplied by the inlet to the throat while simultaneously opening the throat until at least sonic velocity of the fluid stream is attained at the combustion zone, gradually increasing the amount of air supplied by the inlet to the throat and flowing to the combustion zone at such velocity, feeding fuel into such airstream, and restricting the throat adjacent the combustion zone to adjust the velocity, temperature and pressure to the predetermined values required for minimum specific fuel consumption.

3. In a method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet door and an adjustable throat adjacent the combustion zone, and in which as the free airstream strikes the inlet a detached normal shock wave is formed outside said inlet, the steps comprising bypassing a portion of the captured air by opening said adjustable inlet door while regulating said adjustable throat to a substantially completely open position so that said normal shock wave will be caused to move through said inlet and downstream in the duct to the combustion zone, and then closing said adjustable inlet door, feeding fuel into said duct, and regulating said adjustable throat to adjust the thermodynamic conditions to those at which detonation will take place.

4. A method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet door and an adjustable throat adjacent the combustion zone, and in which as the free airstream strikes the inlet a detached normal shock wave is formed outside said inlet, which method comprises bypassing a portion of the cap tured air by opening said adjustable inlet door while regulating said adjustable throat to a substantially completely open position so that said normal shock wave will be caused to move through said inlet and downstream in the duct to the combustion zone, and then closing said adjustable inlet door, feeding fuel into the air stream flowing through the duct, and regulating said adjustable throat so as to adjust the aerothermodynamic conditions in the combustion zone to those at which detonation will take place.

5. A detonation combustion engine comprising: an aerothermodynamic duct having an inlet, a mixing section, a combustion zone having a throat adjacent thereto, and an exhaust nozzle; a door in said duct adjacent said inlet, said door being adjustable between an open position to divert air from said duct to the exterior and a closed position in which no air is diverted; an adjustable ramp immediately upstream of said combustion zone and movable between an open position substantially flush with a wall of the duct and an angular position which reduces the area of the throat adjacent said combustion zone; means for opening said door and moving said ramp to open position during the starting of the detonation combustion engine; and means for feeding fuel into said mixing section.

6. A detonation combustion engine comprising: an aerothermodynamic duct having an inlet, a mixing section, a combustion zone with a throat adjacent thereto, and an exhaust nonle; an adjustable door in said duct adjacent said inlet and movable between a closed position substantially flush with a wall of said duct and an open position which permits part of an incoming airstream to be bypassed; means for injecting fuel into the mixing section of said duct; an adjustable ramp immediately upstream of said combustion zone and movable between an open position substantially flush with a wall of the duct and an angular position which reduces the area of the throat adjacent said combustion zone; and means for opening said door and moving said ramp to open position during the starting of the detonation combustion engine.

7. A method of starting a detonation combustion engine having an aerothermodynamic duct which includes an adjustable inlet and an adjustable throat adjacent the combustion zone and is adapted to receive a hypersonic stream of air during normal operation, which method comprises reducing the amount of air supplied by the inlet to the throat while simultaneously adjusting the throat until a velocity of the fluid stream sufficient to produce a shock wave is attained at the combustion zone, then gradually increasing the amount of air supplied by the inlet to the throat, feeding fuel into the fluid stream, and adjusting the flow conditions in said duct by the regulation of the adjustable throat to obtain the desired velocity, temperature and pressure for detonation in said combustion zone.

8. A detonation combustion engine comprising: an aerothermodynamic duct having an inlet, a mixing section, a combustion zone having a throat adjacent thereto, and an exhaust nozzle; adjustable means adjacent the inlet for changing the amount of air supplied by the inlet to the throat; ramp means upstream of the combustion zone, said ramp means being movable independently of the adjustable means adjacent said inlet to change the area of the throat; means for feeding fuel into said duct; means for moving said adjustable means to reduce the amount of air supplied by the inlet to the throat; and means for moving said ramp means to increase the area of said throat during the starting of the detonation combustion engine.

Patent Citations
Cited PatentFiling datePublication dateApplicantTitle
US2861419 *Feb 13, 1953Nov 25, 1958United Aircraft CorpVariable bleed diffuser
US3040516 *Aug 3, 1959Jun 26, 1962Boeing CoDetonative combustion method and means for ram-jet engine
Non-Patent Citations
Reference
1 *Cushman Scientists Study Mach 7 Ramjet Theory, Aviation Week, Vol. 68, No. 1, Jan. 6, 1958 pages 57 59 and 63
Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US6612522 *Mar 17, 1998Sep 2, 2003Starcraft Boosters, Inc.Flyback booster with removable rocket propulsion module
Classifications
U.S. Classification60/204, 60/767
International ClassificationF02K7/00, F02C7/26, F02K7/10
Cooperative ClassificationF02C7/26, F02K7/10
European ClassificationF02K7/10, F02C7/26