US3719428A

US3719428A - Jet engine for hypersonic intake velocities

Info

Publication number: US3719428A
Application number: US00019418A
Authority: US
Inventors: W Dettmering
Original assignee: Individual
Current assignee: Individual
Priority date: 1969-03-14
Filing date: 1970-03-13
Publication date: 1973-03-06
Anticipated expiration: 1990-03-06
Also published as: GB1273673A; FR2037910A5; DE1913028B1

Abstract

A jet engine, which is suited for hypersonic intake velocities, has a rotating cascade in the intake duct for decelerating the hypersonic flow. The blade angles Beta 1, Beta 2 of the rotating cascade, the circumferential speed U of the rotor and the deceleration Delta w of the relative cascade flow are matched in accordance with cos Beta 2 (U . cos Beta 1)/(U Delta W . cos Beta 1), to ensure an axial direction of the absolute flow at the rotor cascade exit. A stationary vane cascade can be arranged downstream of the rotor cascade, in the case of high blade camber, for further deceleration and axial straightening of the flow. The intake rotor can be coupled to an engine accessory power unit or to a compressor arranged downstream of the intake. Controllable flaps can be provided to route the gas jet leaving the engine combustion chamber either to a turbine arranged downstream for driving the compressor or directly to the engine outlet nozzle. The intake rotor can also be coupled to an acceleration cascade arranged downstream of the combustion chamber; the acceleration cascade can also be preceded by a stationary vane cascade.

Description

United States Patent Dettmering 1 March 6, 1973 [54] JET ENGINE FOR HYPERSONIC FOREIGN PATENTS OR APPLICATIONS INTAKE VELOCITIES 611,447 10/1948 Great Britain ..415/147 Inventor; Wilhelm Heinrich Dettmefing, am 696,007 8/1953 Great Britain ..4l5/l47 Chorusberg 26, Aachen, Germany Primary Examiner-Henry F. Raduazo [22] Fled: March 1970 Attorney-Craig and Antonelli l. N l9 4 8 [2]] App 0 1 57 ABSTRACT [30] Foreign Application Priority Data A jet engine, which is suited for hypersonic intake velocities, has a rotating cascade in the intake duct for March 14, 1969 Germany "P l9 l3 decelerating the hypersonic flow The blade angles B B of the rotating cascade, the circumferential speed U of the rotor and the deceleration Aw of the relative 60/3918 C cascade flow are matched in accordance with cos B [51] Int. Cl. ..F0ld 1/02, F02k 3/00 (U cos BO/(U AW cos 3,), to ensure an axial [58] Field of Search ..60/270, 226,269, 39.18 C; direction of the absolute flow at the rotor cascade 415/181, 147, 143, 79 exit. A stationary vane cascade can be arranged downstream of the rotor cascade, in the case of high [56] Ref n e Cit d blade camber, for further deceleration and axial straightening of the flow. The intake rotor can be cou- UNITED STATES PATENTS pled to an engine accessory power unit or to a com- 2 693,675 11/1954 schafferw 6069M; G pressor arranged downstream of the intake. Controlla- 2:702,985 3/1955 Howell ..415/147 ble flaps can be Pmvided mute the gas let leaving 2,835,470 5/1958 Trowbridge et a1 ..60/39.18 c the engine mbusti0n chamber either a turbine 2,939,017 5/1960 Teague, Jr. et a1. ..60/39.18 c ranged downstream for driving the compressor of 2,952,973 9/1960 Hall et a1. ....60/39.18 C directly to the engine outlet nozzle. The intake rotor 2,945,672 7/1960 Wagner et al..... ....60/39.18C can also be coupled to an acceleration cascade ar- 2,966,028 12/1960 Johnson et al.... ..415/181 ranged downstream of the combustion chamber; the 2,989,843 6/1961 Ferri ..60/226 acceleration cascade can also be preceded by a sta- 3,422,625 1/1969 Harris tionary vane cascade 2,947,139 8/1960 Hausrnann ..415/181 3,442,441 5/1969 Dettmering ..415/181 14 Claims, 5 Drawing Figures (M t-c ("ml/ m ROTOR STATlONARY CASCADE VANE CASCADE c 11 7 I 11 u SHEET 1 or a PATENTED 5 I973 lNVE WILHELM DETT PAIENTEDHM: ems

SHEET 2 [IF 3 ON L) PATENTEDMAR 61% 3,719,428

SHEET 30F 3 ROTOR OUTLET STATIONARY VANE CASCADE INTAKE M l L ROTOR JET ENGINE FOR HYPERSONIC INTAKE VELOCITIES BACKGROUND OF THE INVENTION The present invention relates to a jet engine, and more particularly, to a jet engine suitable for hypersonic intake velocities.

To ensure satisfactory compressor and combustion chamber operation in jet engines with hypersonic intake velocity, it has been necessary, up to now, to decelerate the hypersonic flow in the air air intakes to subsonic velocity. With conventional engines, this has led to compression shocks resulting in static pressure losses and thus to a reduction of total engine efficiency.

The maximum losses occur in a normal compression shock and will increase with the Mach number upstream of the shock. For this reason, deceleration is effected by one or more compression shocks acting in an oblique direction and causing considerably lower static pressure losses or by means of a continuous and thus shock loss-free deceleration achieved with the aid of a wedge-shaped or conoidal bullet. This method, however, necessitates flow-tuming, thereby leading to an increase of the external drag of the engine and thus to net thrust losses. Therefore, the above-described deceleration can only be partially realized and will, at higher flight speeds, e.g., approximately Mach 2.5, not be sufficient to avoid considerable static pressure losses due to an excessive Mach number immediately upstream of the occurring normal compressionshock.

A further deceleration, which causes relatively slight static pressure losses due to flow reversal and simultaneous area restriction, is limited by the set-on of supersonic flow rule, according to which a hypersonic flow in the intake duct can develop only if the mass flow penetrating a normal compression shock in the intake duct does not exceed the capacity of the downstream throat area, otherwise the total deceleration downto subsonic velocity is effected in a normal shock upstream of the duct, resulting in correspondinglyhigh losses. The permissible area restriction and, thus also, the maximum possible deceleration of the supersonic flow becomes higher as the intake velocity C, (Mach number M becomes higher.

SUMMARY OF THE INVENTION It is an aim of the present invention to overcome the problems and disadvantages that have been encountered in the prior art. y

it is a further aim of the present invention to provide 8 maximum reduction of the hypersonic velocity of engine intake flow, combined with minimum losses.

The underlying problems are solved in accordance with the present invention by utilizing the relationship that the minimum-loss deceleration of a hypersonic flow which is made possible by area restriction becomes higher as the Mach number of this flow immediately upstream of the deceleration becomes greater.

The foregoing problems are solved in accordance with the. present invention by arranging a rotating cascade in the intake duct for decelerating the hypersonic flow. Due to the circumferential speed U of this rotating cascade, the relativeflow velocity W, in the cascade channel is increased, since the circumferential velocity U is added, by vector addition, to the absolute intake velocity C which is normal to the cascade plane. Thus, a Mach number M of the cascade flow is obtained and is higher than the Mach number M, of the absolute intake flow. As stated above, it will thus be possible to achieve a higher deceleration, A W=W, -W with a corresponding change in Mach number M=M in the cascade channel while maintaining the supersonic speed, i.e., without a normal compression shock, which will result in a corresponding static pressure increase.

If, by means of suitable blade camber angles, simultaneously and in addition a flow turning against the direction of the circumferential velocity is effected, there will result a considerably reduced absolute exit velocity C, with Mach number M after vector substraction of the circumferential velocity at the cascade channel exit. With an intake velocity of W, having a Mach number M, 2, an arrangement of the intake rotor in accordance with the present invention resulted in a 25 percent increased deceleration as compared to the conventional arrangement without a rotor cascade. Correspondingly, the arrangement in accordance with the present invention resulted in the static pressure ratio in the downstream normal compression shock being increased by 12 percent.

In accordance with an embodiment of the present invention, the blade angles [3 3,, the circumferential speed U of the rotor and the deceleration AW of the relative cascade flow are matched in accordance with the function which will ensure an axial direction of the absolute flow at the rotor cascade exit. This feature prevents any exchange of energy with the rotor so that the latter will not need any driving energy with the exception of the energy to overcome friction.

A further feature of the present invention increases the flow turning effect in the rotor to such a degree, by means of a high blade camber, as to impart a circumferential component to the rotor exit flow. In this connection, a stationary vane cascade is arranged downstream of the rotor cascade for further deceleration and axial straightening of the flow. Due to the more intensive turning of the supersonic flow in the rotor cascade, in addition to the pressure increase by area restriction of the cascade channels as abovedescribed, flow energy is also extracted in the form of shaft power, since an impulse is induced in the rotor in this case. Through this measure, the possible shockfree deceleration AW W, W of the hypersonic flow is still further increased. With an absolute flow inlet Mach number of M 3, an increase in the deceleration of approximately 32 percent is obtained as compared to the version without a rotor cascade. In addition, a corresponding increase of the static pressure ratio in the downstream normal shock of approximately 40% results.

In accordance with a further embodiment of the present invention, the intake rotor can be coupled to an engine accessory power unit for the transfer of power thereto. As a result of this arrangement, the energy extracted from the flow is indirectly returned to the jet engine.

If no accessory drives are to be powered, the intake rotor can be coupled. to a compressor arranged downstream of the intake and thereby transfer the energy extracted from the intake flow to it in accordance with a still further embodiment of the present invention. By including the compressor into the energy transformation taking place in the intake, an excessive compression of the working medium is avoided. Viewed as a system, the cascade rotating in the supersonic range and the compressor coupled to it receive no external energy, thus the total enthalpy between the inlet and exit of the system remains constant. At the high flight Mach numbers under consideration, the pressure increase, which is achieved solely by the lowloss in flow deceleration in the intake rotor-compressor system, is already so high that an additional compression effected by the power transferred from the turbine might lead to an efficiency reduction of the engine, as the temperature rise associated with it will reduce the possible heating margin of the combustion chamber with a given constant maximum temperature.

Another embodiment in accordance with the present invention includes controllable flaps which can be actuated so as to route the gas jet leaving the engine combustion chamber either to a turbine arranged downstream of the combustion chamber fOr driving the compressor or directly to the engine outlet nozzle; in this way, the flow energy extracted by the turbine is controlled, since only very low or even no power is required from the turbine if the compressor is driven by the intake rotor.

In a further development in accordance with the basic idea of the present invention, the energy extracted from the flow within the hypersonic zone at the intake is returned to the engine only downstream of the combustion chamber. Therefore, in an embodiment encompassing this development the intake rotor is coupled to a rotating acceleration cascade arranged downstream of the combustion chamber. The flow acceleration in this cascade is achieved by pressure reduction due to an area increase as well as by work transfer due to intensive turning by the relative flow.

According to still another feature of the present invention, the acceleration cascade downstream of the combustion chamber can be preceded by a vane cascade. The circumferential component of the absolute velocity produced in this cascade suffices to effect the return to an axial direction of the absolute velocity at the rotor outlet. This arrangement is particularly advantageous because it permits control of performance exchange by vane adjustment.

, Still in accordance with the present invention, a vane cascade can be arranged downstream of the rotor cascade located aft of the combustion chamber. This vane cascade serves to straighten the flow from the rotor cascade, which flow has a circumferential component, into an axial direction.

As soon as the gas leaving the combustion chamber is heated to such a degree that it is at least partially ionized, it can also be accelerated by electrical or magnetic fields. The present invention also proposes to produce an electrical or electromagnetic field in the flow ductfof the engine downstream of the combustion chamber, whereby an ionized flow medium is accelerated. The energy required for building up such an acceleration field is produced by a generator driven by the intake rotor.

To permit the use of an engine in accordance with the present invention for missile propulsion in a speed range as wide as possible, the possibility of an adaptation to the respective speed ranges during flight would be desirable. The foregoing is achieved by means of and adjustable rotor and stator blades of the deceleration stage arranged in the intake duct. This relatively simple measure leads to a considerable extension of the speed range in which the missile can operate with advantageous efficiency.

According to a further proposal, the rotor and stator blades of the acceleration cascade arranged downstream of the combustion chamber are made adjustable, thereby similarly extending the speed range in which favorable efficiency can be achieved. In a still further embodiment in accordance with the present invention, at least one coupling can be provided by which the intake rotor is connected to or disconnected from other rotating parts such as, for example, the accessory drive, compressor or acceleration cascade during operation.

In a further embodiment, adjustable flaps are ar ranged in the intake duct for routing the intake flow upon selection either to the intake rotor or through a duct bypassing the intake rotor. This feature is especially. useful in cases where the engine is also intended for subsonic operation.

BRIEF DESCRIPTION OF THE DRAWING These and further aims, features and objects of the present invention will become more apparent from the following description when taken in conjunction with the accompanying drawing which shows, for purposes of illustration only, several embodiments in accordance with the present invention and wherein:

FIG. 1 is a schematic view of an intake of a jetengine in accordance with the present invention with an intake rotor and the associated blade scheme in addition to the velocity triangle diagrams at the cascade inlet and outlet;

FIG. 2 is a schematic arrangement similar to FIG. 1 but with a vane cascade downstream of the intake rotor in addition to the corresponding velocity triangle diagrams;

FIG. 3 is a schematic arrangement of the front section of a jet engine in accordance with the present invention and shows the intake rotor coupled to the first stage of a downstream compressor along with the entropyenthalpy diagram for this jet engine;

FIG. 4 is a schematic view of another embodiment of the jet engine in accordance with the present invention wherein the intake rotor is coupled to an acceleration cascade arranged downstream of the combustion chamber, along with the corresponding entropy-enthalpy diagram; and

FIG. 5 is a schematic view of a still further embodiment in accordance with the present invention, wherein the intake rotor, the compressor and an acceleration cascade arranged downstream of the combustion chamber are connected via a shaft, and adjustable flow control flaps are provided to adapt the engine to the operating condition.

Referring now to the drawings wherein like numerals are used to designate like parts throughout all the views and, in particular to FIG. 1, there is shown one embodiment of the flow intake of a jet engine in accordance with the present invention essentially consisting of a conoidal member 1 arranged at the center of an outer casing 2. To reduce the intake flow losses, a tip 1a of the conoidal member 1 protrudes in a conventional manner substantially beyond front end 2a of the outer casing 2. At the front end 20, an annular flow channel 3 is provided with a rotor cascade 4 which rotates about center line 5. The frontlines of oblique shocks developing at the intake are designated by the numeral 6.

In FIG. 1, there is also shown the blade scheme,of rotor 4 along with the corresponding velocity triangle diagrams upstream and downstream of the rotor 4, wherein:

C absolute velocity U circumferential speed of the rotor W relative flow velocity in the cascade channel and wherein the subscripts used are:

condition of undisturbed flow in front of the conoidal member I= condition upstream of the rotor cascade II condition downstream of the rotor cascade The arrangement shown in FIG. 2 is very similar to that shown in FIG. 1 and described above. However, due to the higher camber of the rotor blade airfoils and the resulting flow velocity downstream of the rotor 4, which has a circumferential component, a vane cascade 7 is arranged downstream. As a result of the camber of vanes 7, the flow velocity C, downstream of the vane cascade is straightened in an axial direction. Auxiliary power means 10 are operatively connected to the intake rotor 4.

The jet engine shown in FIG. 3 has an intake in accordance with the one shown in FIG. 2. Intake rotor 4 is rigidly connected to the first stage 8 of an axial compressor by means of shaft 9. Additional rotors of the axial compressor are designated by the numeral 11, while numeral 12 designates a combustion chamber downstream of the compressor. The circled Roman numerals I, II, III, IV represent the various stages of the jet engine and correspond to the circled Roman numerals in the accompanying entropy-enthalpy diagram.

FIG. 4 shows still a further embodiment of the jet engine in accordance with the present invention which is intended for extremely high speed ranges, e.g., Mach 2.5. Intake rotor 4 is mounted on shaft 9a which can, in turn, be coupled to shaft 9b by means of clutch 13. The clutch 13 can be engaged and disengagedduring operation. Shaft 9b carries an acceleration rotor 14 which is located downstream of the combustion chamber 12. The vanes 7 are arranged downstream of the intake rotor 4.

The acceleration cascade 14 is located downstream of the combustion chamber 12 and is preceded by outlet vane 15 in which a circumferential component is thereby induced into the axially directed flow, The circled Roman numerals are used to designate the conditions of the jet engine as follows:

O Condition of undisturbed flow in front of the conoidal member I= Condition upstream of the intake rotor II Condition downstream of the intake rotor IIl= Condition downstream of the intake vane IV= Condition downstream of the combustion chamber or, alternatively, condition upstream of the outlet vane V Condition upstream of the acceleration vane VI Condition downstream of the acceleration vane VII Flow condition after leaving the jet engine The above conditions have been accordingly entered in the accompanying entropy-enthalpy diagram in FIG. 4.

FIG. 5 shows another embodiment of the jet engine in accordance with the present invention wherein an intake rotor 4, compressor 8, and a turbine 16 which is arranged downstream of the combustion chamber 12 are mounted on common shaft 9. To permit the jet engine to be adapted to various inflow velocity ranges, swivelling flaps 17 are provided in the intake duct which, upon selection, are used to direct the intake air either to the intake rotor 4 or, through a duct, to bypass the intake rotor. Similar flaps 18 are arranged in the outlet duct for routing the flow either to the turbine 16 or to bypass it.

The movement of

flaps

17 and 18 is carried out in a manner such that, with low intake velocities, the intake rotor 4 is covered by flaps l7 and thus the flow is routed through the bypass duct, while flaps 18 open the passage to the turbine 16. At high intake velocities, the pattern is reversed, that is the flow is routed through the intake rotor 4 and the turbine 16 is bypassed. Between these two extreme positions, any desired intermediate position can be selected.

While I have shown and described several embodiments in accordance with the present invention, it is to be understood that the same is susceptible of numerous changes and modifications as will be known to a person skilled in the art, and I, therefore, do not wish to be limited to the details shown and described herein but intend to cover all such changes and modifications as are encompassed by the scope of the present invention.

lclaim:

1. An air intake arrangement for hypersonic air flow duct means for accepting hypersonic air flow entering a jet engine; said arrangement comprising: an air duct means for accepting hypersonic flow, a freely rotatable rotor means adjacent the forward end of said duct means, a shaft supporting said rotor means about a longitudinally extending axis, said rotor means including circumferential spaced blading about said rotor means, said blading extending radially outwardly from adjacent the axis across said duct means, the adjacent blades of said blading defining a plurality of constricting flow paths inclined to the axis of rotation in a downstream direction whereby the air flow is constricted as the air flows downstream along paths rearwardly from between the most forward sections to between the most rearward sections of adjacent blade means, said air flow constriction causing a reduction in the axial velocity of air flow as the hypersonic flow decelerates along the flow paths downstream, the flow velocity in a direction parallel to said flow paths being greater immediately downstream of the most forward sections of said blade means than the flow velocity in the axial direction immediately upstream of said most forward sections, and means for controlling the rotational velocity of said rotor means, said last-mentioned means comprising the geometrical configuration of the blade means whereby static pressure losses are minimized during the deceleration by the constriction of the hypersonic flow through the rotor means.

2. An arrangement according to claim 1, wherein the inlet angle (3 and the exit angle 9,) of the blades of .the rotor means with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration (AW) of the flow through the blades of the rotor means are matched in accordance with the relationship:

COS B2= (U cos ,8,)/(U AW cos (3,)

3. An arrangement according to claim 1, wherein stationary vane cascade means are located downstream of the rotating rotor means for further decelerating the flow and for straightening the flow into a substantially axial direction when the flow leaving the rotating rotor means has a substantially circumferential component.

4. An arrangement according to claim 1, wherein the shaft of the rotor means is operatively connected to the engine for driving an auxiliary power means.

5. An arrangement according to claim 4, wherein the inlet angle (3,) and the exit angle (69 f the blades of the rotor meanS with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration (AW) of the flow through the blades of the rotor means are matched in accordance with the relationship:

cos B2= cos BO/(U AW cos 3,

6. An arrangement according to claim 4, wherein stationary vane cascade means are located downstream of the rotating rotor means for further decelerating the flow and for straightening the flow into a substantially axial direction when the flow leaving the rotating rotor means has a substantially circumferential component.

7. An arrangement according to claim 1 wherein the shaft of the rotor means is connected to a compressor operatively arranged in the engine for supplying power to said compressor.

8. An arrangement according to claim 7, wherein the inlet angle (3,) and the exit angle (B of the blades of the rotor means with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration (AW) of the flow through the blades of the rotor means are matched in accordance with the relationship:

9. An arrangement according to claim 7, wherein stationary vane cascade means are located downstream of the rotating rotor means for further decelerating the flow and for straightening the flow into a substantially axial direction when the flow leaving the rotating rotor means has a substantially circumferential component.

10. An arrangement according to claim 7, wherein the shaft of the rotor means is operatively connected to the engine for driving auxiliary power means.

11. An arrangement according to claim 2, wherein the blades are adjustable.

12. An arrangement according to claim 3, wherein the vane cascade means includes adjustable blades.

13. An arrangement according to claim 1, wherein the velocity of the air leaving the rotor means is supersonic.

14. An arrangement according to claim 1, further comprising auxiliary load power take off means attached to said shaft.

Claims

2. An arrangement according to claim 1, wherein the inlet angle ( Beta 1) and the exit angle ( Beta 2) of the blades of the rotor means with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration ( Delta W) of the flow through the blades of the rotor means are matched in accordance with the relationship: cos Beta 2 (U . cos Beta 1)/(U - Delta W . cos Beta 1)

5. An arrangement according to claim 4, wherein the inlet angle ( Beta 1) and the exit angle ( Beta 2) of the blades of the rotor meanS with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration ( Delta W) of the flow through the blades of the rotor means are matched in accordance with the relationship: cos Beta 2 (U . cos Beta 1)/(U - Delta W . cos Beta 1)

7. An arrangement according to claim 1, wherein the shaft of the rotor means is connected to a compressor operatively arranged in the engine for supplying powEr to said compressor.

8. An arrangement according to claim 7, wherein the inlet angle ( Beta 1) and the exit angle ( Beta 2) of the blades of the rotor means with respect to a plane extending perpendicular to the rotor axis, the circumferential speed (U) and the deceleration ( Delta W) of the flow through the blades of the rotor means are matched in accordance with the relationship: cos Beta 2 (U . cos Beta 1)/(U - Delta W . cos Beta 1)

11. An arrangement according to claim 2, wherein the blades are adjustable.