GB2097863A - Variable air inlet for gas turbine engine - Google Patents

Abstract

A variable air inlet, particularly a two-dimensional oblique-thrust diffuser for a gas-turbine jet propulsion unit for driving a high- performance aircraft, with an upper ramp 1a on a spearhead 1 and with a lower scoop-shaped air interception lip 2 articulated to the front of the air inlet floor 3, and pivotable about a transverse shaft 4, is characterized by the fact that the entire pivoting range of the air interception lip 2 is functionally subdivided into a lower subsonic-transsonic operating zone A and an upper supersonic operating zone B, of a narrower angle. The adjusting angle delta A of the air interception lip 2 in the lower zone A is controlled in accordance with the flight Mach numbers M between 0 and about 1.3 Mach and also the incidence angle alpha of the aircraft, in accordance with the aquation delta A=f (1 DIVIDED M, alpha ), and that the adjusting angle delta B of the air interception lip 2 in the upper zone B is regulated in accordance with a variable required value eta B, as a ratio between the static pressure PSB predominating above, the upper ramp 1a and the overall pressure pt of the outer flow in such a way as to ensure the supply of the air throughput required for the propulsion unit, in order to obtain, in supersonic flight, the optimum thrust. <IMAGE>

Description

SPECIFICATION Air intake for gas turbine driven aircraft This invention relates to an air intake for gas turbine driven aircraft, more particularly to a twodimensional oblique-thrust diffuser intake for gas turbine jet propulsion units for driving highperformance aircraft, such an intake having upper ramps and a lower air interception lip pivoted to the front of the air intake floor and movable about a transverse axis.

The purpose of the air intake of an aircraft is to convert the maximum possible proportion of the kinetic energy of the incident air flow into static pressure energy while reducing the velocity. This recovery of energy may be considerable, particularly at the higher flight Mach numbers. For optimum effect the intake and densified air must be comparatively free of loss and must be conveyed in a homogeneous state to the propulsion unit, in the correct quantity according to the instantaneous performance. Flow losses are mainly caused by air friction, compression shocks and spillage drag. Attention therefore has to be paid to the drag of the external flow, which has to be kept as low as possible. Non-homogeneity in the air is mainly the result of uneven pressures within the flow.If the aforementioned difficulties can be largely avoided, the system comprising the air intake and the gas-turbine propulsion unit will over all power ranges generate thrust with a satisfactory degree of efficiency. This is to a large extent a result of the air intake and propulsion unit interacting in an aerodynamically and functionally stable manner.

At take-off and low flight speeds the propulsion unit needs to be supplied with a considerable volume of air and owing to low pressure at the intake the mechanical air-flow cross-section of the intake at the narrowest point must be made as large as possible. At high supersonic flight speeds, on the other hand, the volume of air through the intake is moderate due to the higher pressures which means that the mechanically narrowest air flow cross section has to be made as small as possible to ensure that in the mouth zone of the air intake the compression shock waves will maintain the required position. High-performance supersonic fighter aircraft are required to be capable of aerial combat engagements in the subsonic range and they have to operate over wide angles of attack.During take-off and at subsonic speeds with high angles of attack and minimum air throughput in the propulsion unit therefore, the air intake lip surface has to be as large as possible but on the other hand for the very small air throughput in the supersonic range, at relatively high air temperatures and under operation at partial load as well as high Mach numbers, the said surface has to be as small as possible. These conflicting requirements cannot be fulfilled by an air intake with a fixed surface, as such an intake in the case of high and extreme angles of attack in flight at subsonic speeds would cause a break-up in the intake flow on a rigid lip leading to compression surging while the low air requirement for the propulsion unit in the supersonic range would produce compression shock wave fluctuations, causing intake hum and instability in the driving system.As a fixed intake has a very limited range of stable operation, highperformance propulsion units are provided with an adjustable air intake which is controlled or regulated in accordance with various flight parameters.

In "Oil Engine and Gas Turbine", Vol. 32, Sept.

1964, a description is given, on pp. 36-39, of a supersonic air intake in the form of a twodimensional adjustable oblique-thrust diffuser with upper middle movable ramps and a bleed device at the inlet end preceding the gas-turbine propulsion unit. In this system the adjustment of the upper movable ramps is effected in accordance with the flight Mach number prevailing while the air bleed flaps are opened when the intake enters the subcritical operating state, that is when the air intake delivers a greater quantity of air than that required by the propulsion unit at that instant.

DE. 23 58 926 also describes an adjustable supersonic air intake with upper movable middle ramps of which the instantaneous position is adjusted in accordance with a measured variable comprising the ratio between the instantaneous static pressure above the movable ramps and the instantaneous over-all pressure of the external air flow. This regulating system ensures that the air intake delivers the air throughput required for the propulsion unit, in order to obtain the optimum thrust at any instant over the entire operating range in supersonic flight.

Both the above constructions have the adjustment of the intake geometry particularly designed for supersonic operation and are not specially suitable for extreme flight conditions in the subsonic range.

DE. 1 066 429 shows a two-dimensional supersonic air intake in the form of a double sided oblique thrust diffuser where the front lips over and under the central cone are made pivotable.

There is no method or apparatus disclosed for controlling or regulating the lower and upper lips in accordance with any particular parameters.

This invention seeks to provide an adjustment method for an air intake and an apparatus for carrying out the method, so that over wide subsonic and supersonic flight ranges and even at extreme flight envelope limits, including take-off, optimum operating conditions are provided by provision of a wide range of operation states between existing aerodynamic stability limits in the intake being hum on the one hand and surging on the other.

According to this invention there is provided an air intake for a high performance aircraft gasturbine jet propulsion unit, the intake including an upper ramp and a lower adjustable air interception lip pivoted to the front of the air intake floor and movable about a transverse axis, the range of adjustment of the lip being divided into a lower velocity subsonic and transsonic range of larger angle extending between a lower mast point having an angle AA-maX, and a central point having an angle 6A,B-0", and a higher velocity supersonic operating range of a smaller angle extending between the central adjusting point and an uppermost adjusting point having an angle bB-m'X, the angle of adjustment SA of the lip in the lower velocity range being controlled in accordance with the flight Mach number M (between 0 and about 1.3 Mach) and the instantaneous attack angle a of the aircraft in accordance with the function

the angle of adjustment SB of the lip in the upper velocity range bB being controlled in accordance with a variable required value 17B formed from the ratio between the instantaneous static pressure Pse above the upper ramp of the intake and the pressure Pt of the incident external airflow whereby the air throughput to the propulsion unit is controlled to obtain in supersonic flight the optimum instantaneous thrust. The middle adjusting point, b, here corresponds or approximately corresponds to the position of an air interception lip of rigid construction, i.e. it here forms the "normal" continuation of the air inlet floor.

The arrangement ensures that in the subsonictransonic flight velocity range the incident air-flow is offered, according to the Mach numbers at which flight is taking place and the angles of attack of the aircraft, a controlled intake profile with a configuration favourable from the point of view of the intake flow so that interruptions in the flow and the resulting compressor surging can be avoided even at extreme angles of attack.

Furthermore, the air intake of the invention operates under supersonic conditions so that the intake cross section is always adapted to the air throughputs required by the propulsion unit in a region somewhat above the critical operating point. In other words, always at a point at which the minimum drag prevails, so that the maximum recovery of pressure or the maximum air throughput will occur.

Preferably for the lower velocity operating range to adjust the lip a control computer means is fed with the instantaneous flight Mach number M and the angle of attack , the computer means having a storage means for required parameters according to Mach number and which have been determined in advance, the parameters providing optimum intake conditions for relatively high and particularly the maximum air throughputs.

Advantageously control of the angle AB of the lip in the higher velocity range is effected through a function of the pressure ratio lie and angle 88 or a value XB proportional thereto, the adjustment parameters 17B-Soll being determined in advance and held in a storage means, adjustment of the lip being effected by comparing the instantaneous required value 17B-Soll with the instantaneous actual value 11B-lst, the lip angle deviation A77B or AXe being determined and reduced to zero by altering the lip angle SB- Preferably the required values 11,-Soll have defined limit reference points w7B-Ref and Xe-Ref, which correspond to defined or selected flight Mach numbers M and flight attack angles a, a change in the operating state or a deviation in the adjustment causing the corresponding or the closest reference point 11,-Ref to be selected.

In a further preferred arrangement lip adjustment is effected by a computer means in accordance with the prevailing flight Mach number M and the flight angle of attack a the reference point X,-Ref of the required value lie Soll being determined and conveyed to a first differentiator which is also fed with the actual value X,-lst or 17B-lst of the lip angle SBT to produce a differential value AxB-Ref which is fed to a function computer to obtain the relevant value A17B-Ref, this value being passed to a second differentiator, a function computer being fed with the instantaneous flight Mach number and lip angle a to determine the reference point 11B-Ref of the required regulating characteristic 17B-Soll which is fed to a second differentiator which subtracts from the value rl,-Ref the value 71B-Ref to produce the required value 17B-Soll of the variable which is fed to a comparator which is fed also with the measured actual value 11,-lst to produce the deviation Afle which is fed to an output computer which determines the regulating signal AXe, passed to a differential value detector which is also fed with the instantaneous value X,-lst of the lip angle s, and from which two values the mechanical adjustment value is calculated.

Advantageously the control computer means and the output computer are connected with a change-over switch which in the lower velocity operating range connects the computer means with a control line and in the higher velocity operating range connects the output computer with said line.

With a supersonic inlet of variable geometry the positions of the movable adjusting means determine the air mass throughputs prevailing at any time, these being the product of the pressure recovery obtained in the air intake and the intake cross section. This value reaches a maximum when the intake is operating in the slightly subcritical range with minimum drag.

In view of the widely varying air throughputs required during operation it is still difficult for the adjusting devices to be always positioned so that the intake will operate in the most satisfactory manner, as the definitive regulating variable stated as "air throughout with maximum pressure recovery" is for all practical purposes no longer measurable in flight and thus has to be discarded as a direct control parameter.

From a practical point of view, therefore, the invention is characterized by the fact that the best possible throughput values are determined with approximate accuracy beforehand, even if at considerable cost as far as measuring technique is concerned, and that on the basis of these values the regulating variable or measured value to which the invention relates is used in flight as the required value in place of the optimum air throughput actually occurring, this being functionally associated with the corresponding lip angle which is set in accordance with the proposed control law. A particular advantage of the invention resides in the fact that the actual values of the regulating variable used can be measured in a simple manner during flight.The same applies to the operation of determining the final co-ordinates for the regulating characteristic required at the time, due to the possibility of accurately measuring the flight Mach numbers and aircraft angle of attack and also to the possibility of determining the exact actual lip angles, from which the required values can be determined for the control variable.

The accompanying drawings show an embodiment of the invention by way of example.

In the drawings: Figure 1 shows a control and adjustment apparatus for an adjustable lip in an air intake and shown schematically, Figure 1 a shows the lip and the adjustment range, Figure 2 shows graphically the adjustment of the lip in the subsonic and transonic speed range A, and Figure 3 shows graphically the adjustment of the lip in the supersonic speed range B.

As shown in Figure 1 , the air intake is constructed as a unilateral oblique-thrust diffuser with a projecting upper rigid lip 1 having an upper ramp 1 a and a lower movable lip 2 which is pivotally mounted on the front of the floor 3 of the intake by a transverse shaft 4. The lip 2 has raised side parts 2a and is scoop-shaped. In the subsonic and transsonic speed ranges of zero to about 1.3 Mach the lip 2 is adjusted over the range A, between a lower adjustment point a, and a middle adjustment point b, equivalent to an angle of 00.

In this case the lip 2 assumes a normal position which is the position of a fixed lip. The angle SA reaches a maximum at the point a. In the supersonic velocity range from about 1.3 Mach upwards the lip 2, over the operating range B, is adjusted between the middle adjustment point b and an upper adjustment point c. The adjusting angle SB being a maximum at the adjustment point c.

The lip 2 is actuated by means of a hydraulic cylinder 5 having a ram 6 coupled to a crank lever 7 which is secured to the lip 2. The ramp 1 a has a gap 8 for air bleed 9, which enters a compartment 10 above the ramp 1 a and exhausts through a grill 11. By means of a pressure sensor 12 the static pressure P55 of this bleed air 9, which varies throughout the entire operating range of the intake is sensed and fed into a transducer 1 3 in which the pneumatic values are converted into electrical signals and passed to a computer 14. By means of a further sensor 1 5 the instantaneous ambient (pilot) pressure Pt of the external flow is determined and also fed to the transducer 13, which passes a signal to the computer 14.The computer determines from the two values p55 and p, the dimensionless ratio 11B-lst, which is conveyed to a comparator 16.

By means of a second pressure sensor 1 7 the static pressure Ps of the external air flow is measured and fed together with the pressure value Pt through the transducer 13, into computer 12 which determines the value PR = Pt/Ps representing the flight Mach number. A further sensor 1 9 determines the instantaneous angle of attack at which the aircraft is flying.

As shown in figure 2, the graph for the lower operating range A of the lip 2 has the aircraft attack angle a along the ordinate and the angle 8A for the lip 2 or the corresponding control signal value XA along the abscissa. The plot lines M1, M2, M3 are also included on the graph, with M1 < M2 < M3. The computer 21 thus determines according to the function

the instantaneous control signal X,-Soll (corresponding to SA~SOII) which is fed via line 26 and Mach changeover switch 24 and a common control regulating line 27, to a comparator 25, which is also fed with the instantaneous angle SA-lst or xA-lst of the lip 2.From these two values AXA is calculated which is the physical measurement in mm for the movement distance of the piston rod 6, or the volumetric quantity of hydraulic fluid fed to the cylinder 5. The cylinder is then supplied with the corresponding quantity of fluid in order to reduce the deviation AXA to zero.

When the lip 2 is being controlled in the range A the Mach value change-over switch 24 occupies a position in which the line 26 is connected with the line 27.

In the graph of Figure 2 the upper aerodynamic limit G3 is that which must not be exceeded to prevent surging due to a sudden interruption of the flow at or in the intake. It may be seen that by moving the lip 2 down as far as possible and despite considerable angles of attack a, acceptable air intake conditions for the propulsion unit are still obtainable. The same applies to relatively high Mach values in the subsonic range, where comparatively high angles of attack a can still be used.

On the other hand the limit G4 shows that with the speeds M1, M2, M3 shown only relatively small angles of attack a are permissible if the lip were not adjustable such as in the case of a rigid lip, as otherwise surging would occur.

The lines M1, M2, M3... are thus required control valves which for relatively high rates of air throughput to the propulsion unit, particularly maximum throughputs, define angles bA in accordance with the instantaneous aircraft attack angle a which are selected by the control apparatus. The control law

indicates that with an increasing angle of attack a the control value XA and thus the lip angle #A increase, so that the lip 2 is moved downwards further With increasing flight Mach numbers M the control magnitude XA and thus the angle #A likewise decreases. The converse is also true.

Care must be taken to ensure that at a particular flight Mach number the maximum permitted angle of attack a is not exceeded or, if exceeded, the fact is signalled.

The graph of Figure 3 for the upper range B of the lip 2 is characterized by the variable ratio lie on the ordinate and the variable value XB, which is proportional to the angle #B of the lip on the abscissa. Of the range of possible air throughput values the air throughput curves have been limited to four, L1 to L4, and L1 > L2 > L3 > L4.The operating range of the propulsion unit and air intake is limited by two aerodynamic ranges G 1 for defining hum and G2 defining surging; Smax denotes the line of maximum thrust of the propulsion unit. lie is in each case a plot of the variable measured value or controlled variable according to which the lip angle #B is set over the operating range B. VB-Soll is the variable which defines the lip angle #B for optimum operation of the air intake and this applies throughout the entire operating range. In other words, VB-Soll and the required lip angle #B-SOll are functionally interrelated and each form a required parameter for the adjustment in the regulating graph ranges.

As mentioned previously the particular angle which is the optimum for the lip 2 at any instant is determined in advance by calculation and by bench and flight tests and the 77B-Soll value functionally related thereto is contained in coded form in the regulator. The 7B-Soll therefore forms the proportional measured value or regulating variable for the angle #B of the lip 2.

As mentioned a number of required regulating characteristics lie are stored in the regulator. For the sake of clarity only one such line 11B-Soll is shown in Figure 3.

The control lines #B-Soll have terminal or reference points which mark the limit coordinates for the values 17B-Ref and X,-Ref. These points add thus the characteristic of the relevant 71B-Soll line are determined from the instantaneous flight Mach number M and the arithmetic value PR = p/pS and the instantaneous angle of incidence a of the aircraft. The diagram of Figure 3 is based on the Mach number M = 1.9 and the angle of attack 3.5 . For all other Flight Mach numbers and aircraft attack angles other 71B-Soll lines apply.

As a result of the regulation according to the invention the two lines SmaX and #B-Soll take advantageously continuous courses.

As may be further seen from Figure 1, the relevant limit coordinate value X,-Ref is calculated in the X#-reference computer 23 in accordance with the instantaneous flight Mach number M and the instantaneous aircraft incidence angle , the same operation taking place in the #B-reference computer 22 for the terminal coordinate value lieRef. In practice the two computers 22 and 23, on the basis of the flight Mach number M and aircraft attack angle a values measured, select the limit coordinates rl,-Ref and Xe-Ref closest to the said measured values.The value X,-Ref determined is fed into a difference computer 28 together with the value X,-lst (instantaneous actual value of the lip angle #B in the operating range B) and the differential value AxB-Ref is calculated from the two values.

This value is then fed to a computer 29 in which the required regulating characteristics 77B-Soll are held. The computer 29, since lie is a function of XB, calculates the values ##B-Ref functionally related thereto. This differential value is conveyed to a required value computer 30 together with the value w7B-Ref and the relevant required regulating point 17B-Soll is determined by subtraction. This value #B-Soll is passed to the comparison computer 16 where it is compared with the #B actual value and the deviation bile determined.

The units 31 and 32 form the dynamic part of the regulator system and in a following output computer 33, using the functionai interrelationship between #B and Xe (cub) #B-Soll is determined from the #B-Soll value. This value is then conveyed through the control signal line 34, the Mach change-over switch 24 and the control signal line 27, to the comparator on 25 which is fed with the instantaneous angle #B or the X,-lst value. From these two values AXB is calculated, which may be the physical measurement in mm for the movement of the piston 6 or the quantity of hydraulic fluid fed to the cylinder 5, as in the case of the control method used in the operating range A. The cylinder is then fed with the required quantity of hydraulic fluid so that the deviation AXB is reduced to zero.

The graph of Figure 3 shows a change in the performance from an operating point p2 to an operating point P3 of lower performance or air throughput. Consequently, during the regulating process, the preceding adjustment angle value XB2 is increased by the differential value Axle, and becomes the value XB3, that is the angle 8e is increased by pivoting the lip 2 upwards as a result of which the air intake mouth becomes smaller.

Claims (7)

1. An air intake for a high performance aircraft gas-turbine jet propulsion unit, the intake including an upper ramp and a lower adjustable air interception lip pivoted to the front of the air intake floor and movable about a transverse axis, the range of adjustment of the lip being divided into a lower velocity subsonic and transsonic range of larger angle extending between a lowermost point having an angle SA-maX, and a central point having an angle #A,B-0 , and a higher velocity supersonic operating range of a smaller angle extending between the central adjusting point and an uppermost adjusting point having an angle gRB-maX, the angle of adjustment #A of the lip in the lower velocity range being controlled in accordance with the flight Mach number M (between 0 and about 1.3 Mach) and the instantaneous attack angle a of the aircraft in accordance with the function

the angle of adjustment #B of the lip in the upper velocity range #B being controlled in accordance with a variable required value #B formed from the ratio between the instantaneous static pressure Pee above the upper ramp of the intake and the pressure Pt of the incident external air flow whereby the air throughput to the propulsion unit is controlled to obtain in supersonic flight the optimum instantaneous thrust.

2. An air intake in accordance with Claim 1, wherein for the lower velocity operating range to adjust the lip a control computer means is fed with the instantaneous flight Mach number M and the angle of attack , the computer means having a storage means for required parameters according to Mach number and which have been determined in advance, the parameters providing optimum intake conditions for relatively high and particularly the maximum air throughputs.

3. An air intake in accordance with Claim 1 or 2, wherein control of the angle #B of the lip in the higher velocity range is effected through a function of the pressure ratio #B, and angle #B, or a value xB proportional thereto, the adjustment parameters 1B-Soll being determined in advance and held in e storage means, adjustment of the lip being effected by comparing the instantaneous required value rl,-Soll with the instantaneous actual value 77B-lst, the lip angle deviation Alive or #XB being determined and reduced to zero by altering the lip angle #B.

4. An air intake in accordance with Claim 3, wherein the required values #B,-Soll have defined limit reference points w1B-Ref and X,-Ref, which correspond to defined or selected flight Mach numbers M and flight attack angles a, a change in the operating state or a deviation in the adjustment causing the corresponding or the closest reference point 71B-Ref to be selected.

5. An air intake in accordance with Claim 4, wherein lip adjustment is effected by a computer means in accordance with the prevailing flight Mach number M and the flight angle of attack a the reference point X,-Ref of the required value rl,-Soll being determined and conveyed to a first differentiator which is also fed with the actual value Xe-lst or #B-ISt of the lip angle #Bt to produce a differential value AxB-Ref which is fed to a function computer to obtain the relevant value A?1B-Ref, this value being passed to a second differentiator, a function computer being fed with the instantaneous flight Mach number and lip angle a to determine the reference point 17B-Ref of the required regulating characteristic #B-Soll which is fed to a second differentiator which subtracts from the value li5-Ref the value rl,-Ref to produce the required value nB-Soll of the variable #e which is fed to a comparator which is fed also with the measured actual value gB-Ist to produce the deviation Alie which is fed to an output computer which determines the regulating signal ##B, passed to a differential value detector which is also fed with the instantaneous value X,-lst of the lip angle #B and from which two values the mechanical adjustment value is calculated.

6. An air intake in accordance with Claim 5, wherein the control computer means and the output computer are connected with a changeover switch which in the lower velocity operating range connects the computer means with a control line and in the higher velocity operating range connects the output computer with said line.

7. An air intake constructed and arranged to function substantially as herein described and with reference to the accompanying drawings.