WO1982000048A1

WO1982000048A1 - Supersonic airplane engine configuration

Info

Publication number: WO1982000048A1
Application number: PCT/US1980/000786
Authority: WO
Inventors: Co Boeing; R Brown
Original assignee: Co Boeing
Priority date: 1980-06-23
Filing date: 1980-06-23
Publication date: 1982-01-07
Also published as: JPS57501096A; EP0054555A4; EP0054555A1

Abstract

A supersonic airplane has a fuselage (20), wings (22) and (24), a vertical stabilizer (34 - 38), a canard (30) and (32) and a pair of turbine engines (44, 46 or 70) mounted in the fuselage (20). Each engine (44, 46 or 70) is of relatively simple de-sign with a fixed inlet, a compressor (80 - 82) and a turbine (84) with a single drive shaft (76) interconnecting the turbine (84) and compressor (80). The engine (44, 46 or 70) can employ either fixed or variable exhaust nozzle geometry (75) depending on whether an afterburner is installed. The engine (70) is point designed to provide maximum design thrust at a given intermediate throttle setting when the airplane is being propelled by the engine (70) at a predetermined supersonic cruise Mach number. Maximum allowable thrust at speeds below the predetermined cruise Mach number is achieved without variable geometry air inlets by maintaining the corrected airflow through the engine (70) at a substantially constant value equivalent to the corrected airflow at the predetermined supersonic cruise Mach number. The engine (70) corrected airflow is maintained at a constant value by monitoring an engine operating parameter indicative of corrected airflow and adjusting the fuel flow to the engine (70) to reduce the actual speed of the compressor (80 - 82) and turbine (84).

Description

SUPERSONIC AIRPLANE ENGINE CONFIGURATION

V

The present invention relates to supersonic airplanes and turbine Engines for such airplanes, and more par- ticularly 'to a single spool turbine engine designed to provide a predetermined amount of thrust at a given throttle setting when propelling an airplane at a pre¬ determined supersonic Mach number, to a method for oper¬ ating such an engine, and to an airplane designed for such an engine.

A conventional turbojet engine is designed to provide maximum thrust at sea level under static operating con¬ ditions, that is, such engines have a sea level static (SLS) design point. The maximum thrust that can be de¬ rived from a conventional turbojet engine at an inter¬ mediate throttle setting (a predetermined high throttle setting without afterburning) is limited by the turbine inlet temperature, other internal engine operating param¬ eters, and limitations imposed by engine construction materials and techniques. The corrected airflow through the conventional turbojet engine with a sea level static design point is maintained at the maximum allowable value from static conditions through transonic flight speeds. Increases in the flight Mach number beyond transonic flight speeds will cause the engine corrected airflow to lapse, that is, drop significantly at an al- most exponential rate, as the flight Mach numbers in¬ crease linearly above the transonic flight regime. The lapse in corrected airflow through the conventional

OMPI turbojet engine begins when the turbine inlet tempera¬ ture rises to its maximum allowable level. Thereafter, increasing compressor inlet temperatures associated with the increased flight Mach number requires a re- duction in fuel flow to the engine, causing the com¬ pressor corrected rotor speed and corrected airflow to decrease while the turbine inlet temperature remains constant at its maximum allowable value.

The significant decrease in corrected airflow through a conventional sea level static (SLS) design engine at supersonic Mach numbers will cause a corresponding drop in the net thrust derivable from the engine at super¬ sonic Mach numbers. In order to obtain the required amount of thrust from a conventional turbojet engine operating at supersonic Mach numbers, the engine must be overdesigned to provide adequate thrust at super¬ sonic flight regimes. Such an overdesigned conventional engine is characteristically large and heavy, both of which characteristics are undesirable. An alternative to overdesigning the engine is to provide the engine with an afterburner, thus providing an augmented power setting to achieve the thrust required for driving an airplane at supersonic Mach numbers. An augmented power setting, however, significantly increases the specific fuel consumption of the engine (pounds of thrust per pound of fuel per hour consumed, hereinafter referred to as SFC) . To achieve fuel efficient supersonic flight speeds, however, it is desirable to operate an engine at ^* an intermediate power setting (that is, a nonaugmented power setting) rather than a partially or fully augmented power setting, thereby lowering the SFC at supersonic speeds. Accordingly, it is the intention to develop a turbojet engine that can provide adequate thrust and associated low specific fuel consumption at supersonic flight speeds. Another broad object of the present invention is to provide an airplane with inlet and exhaust nozzle configurations that are relatively simple in construc¬ tion and require no variable geometry engine inlet structure or variable geometry nozzle structure for intermediate throttle settings.

Summary of the Invention

The present invention provides an improved turbojet en¬ gine that is point designed to provide a predetermined amount of thrust at a given throttle setting when pro¬ pelling an airplane at a predetermined supersonic flight Mach number. The improvement includes a control means for maintaining the corrected airflow through said en¬ gine at a substantially constant value at flight Mach numbers below the predetermined supersonic flight Mach number. Preferably, the turbine engine of the present invention is of the single spool type, including a compressor, a turbine and a single drive shaft inter¬ connecting the turbine and compressor. More preferably, the turbine engine has associated with it a fixed geometry inlet for ducting air to the compressor. The engine is preferably point designed to economically operate at an intermediate throttle setting at super¬ sonic flight Mach numbers in the range of from 2.0 to about 2.7. An afterburner and variable geometry ex¬ haust nozzle can be employed with the turbine engine of the present invention to provide augmented transonic thrust to shorten the acceleration time of the airplane

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OV.PI through the transonic flight regime.

An additional aspect of the present invention includes a method for controlling the turbine engine described above by sensing an engine operating parameter such as the turbine or compressor pressure ratio to provide a signal representative of the corrected airflow through the engine. This signal is then compared with a second signal representative of a predetermined maximum value of the engine operating parameter. A control signal, derived from and representative of the difference be¬ tween the first and second signals, is then employed to adjust the operating condition of the engine to maintain the corrected airflow through the engine constant at flight Mach numbers below the predetermined supersonic flight Mach number. For engines of the present inven¬ tion that are point designed to operate, at supersonic flight Mach numbers on the order of 2.7, the angle of the turbine rotor blades of the engine can be increased relative to the stator blades at.flight Mach numbers below the predetermined supersonic flight Mach number.

In a further aspect of the present invention, an airplane having a wing, a vertical stabilizer, a canard and a turbine engine mounted in the fuselage is provided. The turbine engine for the airplane is point designed to- provide maximum thrust at a given intermediate throttle setting when the airplane is being propelled by the en¬ gine at a predetermined supersonic flight Mach number. The engine includes control means for maintaining the corrected airflow through the engine at a substantially constant value at flight Mach numbers below the pre¬ determined supersonic flight Mach number. The airplane has an engine air inlet positioned adjacent the rear¬ ward portion of the canard for ducting air to the tur¬ bine engine and a nozzle located at the aft portion of the fuselage for directing exhaust gasses rearwardly from the engine. Preferably, the canard is located on the fuselage above and forwardly from the air inlet and is so positioned relative to the inlet to provide pre- compression of the air flowing under the canard and in¬ to the air inlet while the airplane is in flight. If maximum acceleration is desired, the airplane and en¬ gine can also employ an after burner and a variable geometry exhaust nozzle to increase the available thrust in the transonic flight regime.

Brief Description of the Drawings

A better,understanding of the present invention can be derived by reading the ensuing specification in conjunc¬ tion with the accompanying drawings, wherein:

FIGURE 1 is a plan view of an airplane configured in accordance with the present invention in which the point design turbojet engine of the present invention is em¬ ployed;

FIGURE 2 is an elevation view of the airplane of FIGURE 1;

FIGURE 3 is a front view of the airplane of FIGURE 1;

FIGURE 4 is a simplified view in partial longitudinal section of an engine designed in accordance with the

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OΛΪPI present invention shown with an accompanying block diagram illustrating engine operating parameter sensors, comparators, and fuel flow regulation apparatus for maintaining corrected airflow through the engine at a constant value for all flight Mach numbers;

FIGURE 5 is a graphic comparison of the engine corrected airflow in pounds per second at selected flight alti¬ tudes through the point design turbojet engine of the present invention and a conventional_,,SLS turbojet engine versus flight Mach number;

FIGURE 6 is a graphic comparison of the maximum turbine temperature in degrees Rankine at selected flight alti¬ tudes of the point design turbojet engine of the pres¬ ent invention and a conventional SLS turbojet engine versus increasing flight Mach number;

FIGURE 7 is a graphic comparison of the uninstalled net thrust in thousands of pounds at selected flight alti¬ tudes for the point design turbojet engine of the pres¬ ent invention and a conventional SLS turbojet engine versus increasing flight Mach number;

FIGURE 8 is a graphic comparison of the uninstalled specific fuel consumption in pounds of thrust per pound of fuel consumed per hour at selected fl ght altitudes for the point design turbojet engine of the present in- vention and a conventional SLS turbojet engine versus increasing flight Mach number;

FIGURE 9 is a graphic comparison of uninstalled aug¬ mented thrust in thousands of pounds at selected flight

BU K tA G.MPI altitudes for the point design turbojet engine of the present invention and a conventional SLS turbojet en¬ gine versus increasing flight Mach number;

FIGURE 10 is a simplified view in partial longitudinal section of another embodiment of the point design turbo¬ jet engine of the present invention having variable angle turbine blades and including a schematic block diagram of engine parameter sensors, comparators and fuel flow regulation apparatus for maintaining corrected airflow through the engine at a constant value for all flight Mach numbers; and

FIGURE 11 is a simplified sectional view taken along section 11-11 of FIGURE 10 showing the variable position turbine blades in conjunction with a black diagram illustrating a flight condition sensor, a comparator, and apparatus for varying the position of the turbine blades in response to a control signal from the com¬ parator.

Detailed Description

Certain terms used herein are of a specialized nature but are known to those of ordinary skill in engine and airplane technology. However, to facilitate a ready and complete understanding of the invention, definitions for certain of these terms are set forth below.

The term "turbojet engine" is intended to encompass all jet engines employing a compressor, a combustor and a turbine for generating a high energy exhaust stream that is directed away from the engine to yield a reactive thrust component on the engine and consequently on an airplane in which the engine is mounted.

The phrase "intermediate throttle setting" is utilized to describe a throttle setting for a turbojet engine which causes the engine to generate a predetermined amount of thrust at or near the upper limit of the en¬ gine's design capability without the use of afterburn¬ ing apparatus. The term "intermediate thrust" is used to describe the thrust derived from a turbojet engine at an intermediate throttle setting. The term "augment¬ ed throttle setting" is utilized to describe a throttle setting for a turbojet engine having afterburning appa¬ ratus which causes the engine to generate the thrust available from the engine at an intermediate throttle setting plus the thrust available through the operation of the afterburning apparatus. The term "augmented thrust" is used to describe the thrust available at an augmented throttle setting.

The term "single spool" is utilized to define that por¬ tion of the turbojet engine consisting of a compressor, a turbine and a single drive shaft interconnecting the turbine with the compressor. A single spool engine may have one or more stages in each compressor and the tur- bine.

The term "engine Mach number" refers to the Mach number of the air flowing through or past any given point in the engine downstream of the compressor.

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O.V.PI The term "flight Mach number" refers to the flight speed in Mach number of a jet propelled airplane.

The term "inlet Mach number" is utilized to refer to the Mach number of the air flowing past the face of the in- let duct to the compressor.

The term "engine corrected airflow" or "corrected air¬ flow" refers to the temperature and pressure corrected mass flow rate of air through the engine. Engine cor¬ rected airflow is defined as

w„

wherein is the actual mass flow rate of air in pounds per second at the inlet face of the engine,

θ is the total temperature at the face of the engine in degrees Rankine divided by 519.4° Rankine, and

δ₂ is the total pressure at the face of the engine in pounds per square foot divided by 2116 pounds per square foot.

The term "inlet geometry refers to the structure con¬ figuration of the inlet duct leading to the compressor of a turbine engine. The phrase "variable inlet geomet¬ ry" refers to inlet structure associated with the inlet ducts that is movable to vary the cross sectional con¬ figuration of the inlet duct. In contrast "fixed inlet geometry" refers to inlet structure that is configured in a permanent or fixed relationship and contains no

-_BURtA movable structure to vary the cross sectional geometry of the inlet duct. As a corollary, "fixed nozzle geom¬ etry" refers to an exhaust nozzle structure having a fixed configuration, while "variable nozzle geometry" refers to exhaust nozzles having movable walls to vary the cross sectional configuration of the nozzle during use.

Referring to FIGURES 1, 2 and 3, one airplane with which the novel engine of the present invention can be employed generally has an elongated fuselage 20 and a delta wing configuration including a left wing 22 and a right wing 24. Segmented elevons 26 and 28 are mounted on the aft portion of the left and right wings 22 and 24, respec¬ tively, and integrate the functions of conventional elevators and flaps. The elevons are provided in four segments, two segments on each wing, to permit safe flight if any one of the elevon segments should fail. Left and right canard segments 30 and 32 respectively are mounted on the upper, forward half of the fuselage and terminate in trailing edges adjacent the root and above the leading edge of the delta wings 22 and 24. In the embodiment shown, the canard has no aerodynamic control sufaces, but can employ such surfaces if de¬ sired. The canard has an aerodynamically curved, posi- tive camber upper surface and a substantially flat bot¬ tom with a slight under camber. The canard has a slight positive angle of incidence relative to the delta wings. The preferred angle of incidence is a positive 2°. The preferred angle of incidence for the canard permits the wing trailing edge controls to always be in a zero trim condition or slightly down when the wing is at a cruise angle of attack. The canard also fucntions to move the aerodynamic center of the airplane forwardly of that which would result if no canard were used, resulting in a long tail arm to the trailing edge wing controls. The long tail arm allows the pilot to maneuver the air- plane, especially in the pitch moment, with smaller movements in the elevons.

A dorsal tail surface 34 having an aerodynamic cross section is mounted on the aft portion of the fuselage and includes a rudder 36 for yaw control. A ventral tail surface 38 is also mounted on the aft portion of the fuselage to provide additional directional stability for high maneuverability at supersonic Mach numbers. The ventral tail surface 38 can be hinged along a hinge line 40 so that the lowermost segment 42 of the ventral tail surface can be swung upwardly to provide adequate ground clearance upon takeoff and landing.

Two engines 44 and 46 are mounted in the aft portion of the fuselage: The engines 44 and 46 will be de¬ scribed in substantially more detail later in the speci- fication. However, it is to be understood that the en¬ gines are point designed for maximum thrust at an inter¬ mediate throttle setting at flight Mach numbers between 2.0 and 2.7. For the airplane shown, it is preferred that the engines be point designed for maximum thrust at an intermediate throttle setting for flight Mach numbers on the order of 2.7. The engines terminate in exhaust nozzles 48 and 50, to be described in more de¬ tail below, which direct the high energy exhaust rear- wardly from the fuselage 20 of the airplane. The en- gines and the nozzles are oriented so that the exhaust stream emanating from the nozzles is directed at a small angle inwardly toward the longitudinal center line of the airplane. This orientation of the exhaust stream reduces the overall drag characteristics of the airplane and thus increases the airplane efficiency. Combustion air is fed to the inlet faces 52 and 54 of the engines 44 and 46, respectively, through inlet ducts 56 and 58, respectively. The inlet ducts 56 and 58 open forwardly in rectangular openings 60 and 62, the planes of which are oriented substantially ortho¬ gonally to the longitudinal center line of the fuselage 20. The inle ducts 56 and 58 extend rearwardly from the rectangular openings 60 and 62 and transition from a rectangular cross section at the forward end of the ducts into a circular cross section at the rearward end of the ducts equivalent to that of the inlet faces 52 and 54 of the engines. The inlet openings 60 and 62 to the inlet ducts are spaced outwardly from the sides - of the forward portion of the fuselage 20. Entrance ramps 64 and 66 extend forwardly and inwardly relative to the fuselage from the inner vertical edges of the duct inlets and fair into the fuselage below the mid- portion of the canard segments 30 and 32, respectively. The inlet openings 60 and 62 are positioned vertically immediately forwardly of the trailing edge of the canard segments 30 and 32 and have their upper edge positioned adjacent the trailing edge of the canard segments, leaving only a small gap between the upper edge of the inlet openings and the trailing edge of the canard sur- faces. Since the vertical ramps 64 and 66.are completely in the canard flow field, the orientation and relation¬ ship of the canard segments with respect to the inlet

OM ramps protect the inlet from all yaw and pitch cross flows during flight of the airplane. At cruise angles of attack, the canard flow field provides precompression of the air entering the inlet, lowering the inlet Mach number below that of the flight Mach number. Thus the inlet can be point designed for an inlet Mach number lower than that which would otherwise be required if the inlet geometry just described were not employed. In the preferred airplane and engine configuration illustrated in FIGURES 1 through 3, the entrance ramps 64 and 66, the inlet openings 60 and 62 and the inlet ducts 56. and 58 all have fixed geometry, providing a relatively simple inlet structure in the fuselage and eliminating the need for extra controls, variable geometry inlet surfaces and the associated weight-increasing variable geometry sur¬ face actuating apparatus.

Refer_/ring still to FIGURES 1 through 3, and especially to the latter, the extremely large capture area for the engine, including the entrance ramps 64 and 66 and the _. inlet openings 60 and 62, provide the airplane with a very small frontal area, which comprises only the nose portion and that portion of the fuselage rearwardly of the nose that extends vertically into the airstream. The small frontal area results in an airplane having very low drag, which in turn results in a more fuel efficient airplane because relatively smaller engines can be utilized than would otherwise be required.

Referring now to FIGURE 4, one embodiment of the point designed turbojet engine of the present invention is illustrated. The engine 70 includes a housing 72, inlet

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OMPI face 74, and a variable geometry nozzle 75. The engine has a single spool, including a single drive shaft 76 mounted for rotation in bearings 78, that directly couples in driving relationship the first section of the compressor, generally designated 80, and the second section of the compressor, generally designated 82, with the turbine, generally designated 84. The first section 80 of the compressor consists of three sets of axially spaced, stationary guide vanes 80a, 80b and 80c each followed by the rotating compressor blades 80d, 80e and 80f coupled to the drive shaft 76 via compressor wheels 80g, 80h and 80i. Similarly, the second section 82 of the compressor comprises four sets of axially spaced stationary guide vanes 82a, 82b 82c and 82d preceded by four sets of rotating blades 82e, 82f, 82g and 82h in turn coupled to thd drive shaft 76 by axially spaced compressor wheels 82i, 82j, 82k and 821. An annular combustor generally designated 86 is located aft-of the compressor. Fuel is sprayed into the combustor 86 at a location aft of the compressor by spray bar 90. As will be understood by one of ordinary skill in the art, more than one spray bar 90 is normally employed in the engine to supply fuel to the combustor, although only one is illustrated for purposes of simplifying the ex- planation of the engine. The turbine 84 is situated aft of the combustor and receives the pressurized products of combustion from the combustor 86. The products of combustion are expanded through the turbine 84. The turbine 84 comprises three axially spaced sets of stator blades 84a, 84b and 84c, sandwiched between which are two sets of rotor blades 84d and 84e. The rotor blades 84d and 84e are coupled via turbine wheels 84f and 84g to the drive shaft 76. As the products of combustion from the combustor 86 are expanded through the turbine 84 in a conventional manner, the turbine rotor, consisting of the rotor blades and rotor wheels, turns the drive shaft, which in turn drives the compres- sor to compress combustion air entering through the in¬ let face 74 of the engine.

To achieve the engine corrected airflow required to pro¬ vide the high intermediate thrust levels at and above the Mach 2.0 design point, the compressor blade angles must be about 20% to 40% greater than the compressor blade angles employed in conventional engines. The com¬ pressor blade angles in the conventional- turbojet are on the order of 22° (relative to a plane orthogonal to the centerline of the engine) , while the required compressor blade angles for the point design engine must be on the order of 26°, or greater if desired. The higher blade angle will provide the required axial flow of combustion air to the combustcr at cruise conditions.

Optionally, the engine can include conventionally de- signed afterburning apparatus, simply shown in FIGURE 4 as a spray bar 90 located in the throat of the nozzle aft of the turbine 84. As will be understood by one of ordinary skill in the art, a plurality of spray bars similar to 90 are normally employed to provide adequate fuel and even combustion in the afterburner. Fuel is supplied to the afterburner spray bar 90 via line 91 from a fuel source (not shown) . The fuel supply through line 91 is regulated by conventional valving apparatus responsive to pilot controlled throttles in a conven- tional manner. If the optional afterburner apparatus is employed, it is preferred to also employ the variable geometry nozzle 75 that adj sts the nozzle throat area between a first, restricted position for intermediate throttle settings (as shown in solid outline) and a sec- ond open position for augmented throttle settings. When the afterburning apparatus is in operation, the throat area of the nozzle is enlarged to increase the exit area of the nozzle by moving the variable nozzle structure outwardly to the position shown in dot-dash outline 75a. The exit area of the nozzle is increased to relieve back pressure on the turbine at augmented throttle settings and to appropriately match the flow rate and pressure in the nozzle with the corresponding speed and temperature of the turbine at both intermediate and augmented throt- tie settings. The control and actuation apparatus for the variable geometry nozzle structure are not shown or described as any of a variety of conventional variable geometry nozzles can be employed with the engine 70,

The engine illustrated in FIGURE 4 is point designed at the cruise condition to provide the desired high inter¬ mediate thrust levels at supersonic Mach numbers on the order of 2.0. The engine is point designed for maximum efficiency and thrust at flight Mach numbers of Mach 2.0 without compromise to off design operating conditions, that is, without compromise to performance at flight

Mach numbers below 2.0. Point designing the engine for maximum thrust at Mach 2.0 requires the engine config¬ uration, including the inlet face size, compressor con¬ struction, combustor construction, turbine construction and nozzle construction, to be optimized so that the engine maximum turbine inlet temperature, compressor ratio, corrected airflow, corrected rotor speed and physical rotor speed are all designed to achieve a maxi¬ mum at the Mach 2.0 ^'.cruise condition. The engine off design physical speed is controlled as a function of compressor match point at intermediate power settings for all flight speeds. That is, the off design physi¬ cal speed of the rotor is varied so that the engine Mach number and the engine corrected airflow is maintained constant for all flight speeds below the point design conditions of Mach 2.0. Point designing and engine for maximum thrust at an intermediate throttle setting for flight speeds of Mach 2.0 allows fixed geometry inlet structure to be utilized and provides a near optimum match of the fixed geometry inlet structure to engine airflow demand at the supersonic flight speed design point. To achieve the near optimum match of the engine airflow demand to the inlet structure of a conventional SLS turbojet engine would require the use of variable geometry inlet structure. The high corrected airflow that can be employed at the point design cruise con¬ ditions with the engine of the present invention allows utilization of a larger nozzle exit area relative to the conventional SLS engines at high Mach numbers. The en- gine of the present invention can achieve installed intermediate thrust increases at cruise conditions on the order of 50% over equally sized, fixed geometry inlet, conventional SLS turbojet engines at Mach flight numbers of 2.0

The point design engine of the present invention, how¬ ever, when designed to provide adequate thrust for a

OMPI given airplane configuration at the point design con¬ ditions (for the engine illustrated in FIGURE 4, flight Mach numbers of on the order of 2.0) will suffer an inter¬ mediate throttle setting thrust lapse with decreasing flight Mach numbers because the turbine temperature, i.e. , the total temperature at the face of the turbine, must be reduced at flight Mach numbers below the super¬ sonic cruise condition to avoid exceeding the upper limit of the compressor corrected rotor speed. If the total temperature (or the engine corrected airflow) were al¬ lowed to exceed the predetermined maximum, the physical speed of the rotor would exceed the structural limita¬ tions of the rotor and could cause physical damage to or disintegration of the engine. The transonic intermedi- ate thrust lapse may, under certain conditions, be too severe to provide adequate airplane transonic accelera¬ tion. That is, under transonic flight conditions, the engine of the present invention may provide inadequate thrust to accelerate the airplane in which it is in- stalled through the transonic flight regime in a time interval acceptable to the user of the airplane. For this reason, it is desirable and preferred to install afterburning apparatus to augment the net thrust de¬ rivable from the engine in the transonic flight regime. Thus, an augmented throttle setting can be utilized with the point design engine of the present invention to offer thrust levels for acceleration through the transonic flight regime comparable to those derived from conven¬ tional SLS turbojet engines of equivalent size and weight, while retaining the high intermediate thrust levels and economy of the point design engine for the supersonic or point design supersonic speeds. The engine corrected airflow is maintained at a constant level for all flight Mach numbers below the point design Mach number by varying the fuel input to the engine to decrease the physical rotor speed at flight Mach number below the point design Mach number. In the embodiment of the invention illustrated in FIGURE 4, an engine op¬ erating parameter than can be corollated to and is rep¬ resentative of the engine corrected airflow is sensed. The sensed engine operating parameter is compared with a reference signal indicative of the maximum and desired engine corrected airflow at an intermediate throttle setting to provide a control signal that in turn is rep¬ resentative of the difference between the reference sig¬ nal and the sensed engine operating parameter. If the difference in the sensed and reference signals indicates that the engine corrected airflow is rising above the maximum point design corrected airflow, the control sig¬ nal is utilized to adjust the physical rotor speed down¬ wardly to maintain the engine corrected airflow.at a constant, value corresponding to that indicated by the reference signal. For example, if the engine is oper¬ ating at an intermediate throttle setting at a flight Mach number of 2.0 and the airplane is slowed down to a flight Mach number below 2.0, the output signal de- rived from the sensed engine operating parameter and the reference signal will indicate that the engine cor¬ rected airflow is beginning to exceed or is exceeding the desired maximum, requiring a reduction in the phys¬ ical rotor speed. The control signal is then utilized to reduce the fuel flow to the combustor of the engine and thereby reduce the turbine inlet temperature and the physical rotor speed. Once the physical rotor speed is reduced sufficiently, the sensed operating parameter will match with the reference sign indicating that the engine corrected airflow is back to its predetermined maximum for the given flight Mach number.

By way of another example, when the engine is acceler¬ ating the airplane through the transonic flight regime at an intermediate throttle setting that would normally provide maximum thrust at the point design conditions, the engine corrected airflow will tend to fall below the maximum engine corrected airflow at the design point. Under these conditions, the sensed operating parameter would provide an indication that the engine corrected airflow is tending to fall below or is below the pre- determined maximum. Thus, the sensed engine parameter, when compared with the reference signal, will provide an output signal which in turn can be utilized to in¬ crease the fuel input to the engine and control the physical rotor speed at a constant value. Still refer- ring to FIGURE 4, one engine operating parameter that can be sensed to provide an indication of the engine corrected airflow through the engine is the turbine pressure ratio, that is, the ratio of the pressure at the turbine inlet to the pressure at the turbine outlet. The turbine pressure ratio can be sensed by positioning a pressure sensor 92 at the turbine inlet face to sense the pressure at the turbine face. A second pressure sensor 94 can be positioned immediately aft of the turbine to sense the pressure at the rear of the turbine. The pressure related signals provided by the pressure sensors 92 and 94 are transmitted via lines 96 and 98

A to a comparator 100. The comparator 100 conditions the inlet and outlet pressure signals provided by lines 96 and 98 to provide an output signal that is representa¬ tive of the actual engine corrected airflow through the engine based upon the turbine pressure ratio.

The output signal representative of the actual corrected airflow through the engine is transmitted through line 102 to a second comparator 104. The reference signal, which is indicative of the maximum point design engine corrected airflow, is input to the comparator 104 via line 105 or may be embodied in a memory in the compara¬ tor 104. The actual engine corrected airflow is then compared with the maximum engine corrected air low ref¬ erence signal in comparator 104 to provide a control signal representing the difference between the actual and reference engine corrected airflow signals. The control signal can be utilized to control apparatus for adjusting the turbine inlet temperature to prevent the actual engine corrected airflow from exceeding the axi- mum point design (or reference) engine corrected air¬ flow. For example, if the actual engine corrected air¬ flow should begin to exceed the reference engine cor¬ rected airflow, the differential output signal from the comparator 104 is transmitted to a servocontrol 106 via line 108 to energize the servocontrol. The servocontrol 106 in turn is coupled to a proportional fuel flow con¬ trol valve 110 interposed in the fuel supply line 112 to the spray bar 90 in the combustor 86 of the engine. The fuel flow control valve is positioned in the fuel line 112 so as to override the manual throttle setting

1JUREA

OMPI to the primary fuel flow control valve 116. Thus if the control signal in line 108 indicates that the en¬ gine corrected airflow is beginning to exceed the ref¬ erence engine corrected airflow, the control signal will energize the servocontrol 106 to partially close the flow control valve 110, thus reducing the quantity of fuel being introduced to the combustor 86. As a con¬ sequence, the turbine inlet temperature will be lowered, the actual rotor speed will be reduced and the turbine pressure ratio will drop. As this occurs, the pressure sensors 92 and 94 will sense the reduction in turbine pressure ratio, transmit it to comparator 100, from which the signal representative of the reduced actual engine corrected airflow will be transmitted to the second comparator 104. When the actual engine corrected airflow drops sufficiently to match the reference en¬ gine corrected airflow, the control signal from the comparator will de-energize the servocontrol 106 and thus the override flow control valve 110 at a position that will maintain the fuel flow to the combustor at the desired rate. Vice versa, as the flight Mach num¬ ber of the airplane in which the engine is mounted in¬ creases, thus allowing the turbine inlet temperature to be increased, the turbine pressure ratio sensors 92 and 94 will provice a signal via the first comparator 100 indicating the actual engine corrected airflow is at a value below that of the reference engine corrected air¬ flow. The control signal from the second comparator 104 will then energize the servocontrol 106 to open the override control valve 110 to increase the fuel flow to the engine and thus increase the turbine inlet tempera¬ ture and actual engine corrected airflow up to the de¬ sired maximum value.

"BURE The control system for maintaining the engine corrected airflow constant at the maximum desired value at all flight Mach numbers below the point design Mach number is an override control system that overrides the manual throttle setting chosen by the pilot. In a conventional manner, fuel is supplied to the engine via the fuel line 112 from a fuel source 114 including a fuel storage tank, fuel pumps, and fuel metering devices. A primary fuel control valve 116 is interposed in the fuel line 112 downstream from the fuel source 114. The primary fuel control valve is controlled by a setting of the manual throttle 118. Thus, if at takeoff the pilot sets the manual throttle 118 wide open at an intermediate setting, the fuel control valve 116 will be open to its maximum extent, allowing fuel to flow to the override fuel con¬ trol valve 110 for maximum intermediate thrust. How¬ ever, as the engine speed increases, the turbine pres¬ sure ratio will rise and will provide a.signal to the override control valve 110 to meter the fuel to the spray bar 90 in order to maintain the engine corrected airflow at the maximum allowable value as the airplane accelerates from static ground conditions to the design point flight Mach number.

The graphs of FIGURES 5 through 8 comparatively illus- trate several operating characteristics of the point de¬ sign engine of the present invention as contrasted with a similarly sized conventional SLS turbojet engine. For purposes of the comparison, certain design param¬ eters are utilized to define the point design engine to the present invention and the conventional SLS turbojet engine cycles. These engine design parameters are set forth in TABLE 1 below: TABLE 1 ENGINE DESIGN PARAMETERS DESIGN PARAMETERS CONVENTIONAL POINT DESIGN

Design Point SLS¹ M=2.0, 60,000 F

Corrected Airflow (lb/sec) 400 400

Compressor Pressure Ratio 10 _, 10

Compressor Efficiency (%) 83 83

Burner Εxit Temp. (°F) 3,000 3,000

Turbine Efficiency (%) 87 87

Cooling Flows

Percent Engine Airflow 13.8 13.8

Pressure Losses - P/P

Main Burner .06 .06

Turbine Exit .02 .02

AB Dry⁴ .03 .03

Afterburner

Temp Max (°F) 3340 3340.

Efficiency (%) 89 89

Entrance Mach No. .26 .26

Inlet Recovery 1. .925

Nozzle CV⁵ .985 .985

Nozzle CD⁶ 1.0 1.0

Footnotes to Table 1:

1. The conventional engine is designed to provide maximum thrust at an intermediate throttle setting at sea level static (SLS) conditions.

2. The point design engine is designed to provide maximum thrust at a flight Mach number of 2.0 at an alti¬ tude of 60,000 feet.

3. The engine airflow is pressure and temperature corrected to standard conditions according to the formula set forth in the first part of the Detailed Description.

4. AB is an abbreviation for afterburner.

5. CV is the velocity coefficient.

6. CO is the discharge coefficient.

As can be seen from the design parameters set forth in Table 1, the conventional engine is designed for maximum thrust at an intermediate throttle setting at sea level static conditions while the point design engine is de- signed for maximum thrust at an intermediate throttle setting at a flight Mach number of 2.0 and an altitude of 60,000 feet. The corrected airflow, the compressor pressure ratio, compressor efficiency, burner exit temp¬ erature, turbine efficiency, cooling flows, pressure losses and afterburner design parameters are all assumed to be equal. The inlet recovery for the conventional engine is assumed to be 1 while that of the point design engine is assumed to be 0.925. Both the velocity co¬ efficient and the nozzle coefficient are assumed to be equal for both engines.

With the foregoing design parameters, the estimated weights and dimensions for the point design engine of the present invention and for the conventional engine are substantially the same, although the diameter and length of the nozzle are slightly larger for the point design engine to reflect appropriate sizing of the aug- enter duct for increased corrected airflow at a cruise flight Mach number of 2..0. Because of the slightly larger nozzle, the weight of the point design engine will be increased by a few percent. For purposes of the comparison, a fixed geometry half round inlet was selected to illustrate engine installed performance. The inlet selected has a single, fixed 25° cone with no bleed and no variable geometry. The inlet design pro- vides minimum inlet complexity, an advantage of the point design engine, and simulates the inlet design de¬ scribed in conjunction with the airplane shown in FIG¬ URES 1 through 3.

-_BUK AL

OMPI With the foregoing engine design point parameters and structural assumptions, the off design variation in the engine operating parameters at intermediate thrust set¬ tings for the point design engine dictate that the point design engine be operated at a constant engine corrected airflow, engine face Mach number, compressor pressure ratio, and corrected rotor speed. (Corrected rotor speed is the actual rotor speed divided by a temperature correction factor / Q , wherein/ Q is equal to the total inlet temperature.at the face of the engine divided by 519.4° Rankine.) The turbine temperature of the point design engine varies proportionally to the ratio of the total inlet temperature at the face of the engine divided by the maximum design temperature at the face of the en- gine. The rotor speed is thus variable with the total inlet temperature. For both the point design engine and the conventional engine, the nozzle throat area remains constant, that is, no variable geometry is employed at intermediate thrust settings. For the conventional en- g ne, engine corrected airflow, engine face Mach number, compressor pressure ratio compressor efficiency and cor¬ rected rotor speed are all constant at values of Θ₂ below 1, and all decrease with increasing total temperature at the face of the engine at values of e₂ above 1. The tur- bine temperature for the conventional engine is constant at values of Q₂ above 1, and at values of Θ. below 1 var¬ ies proportionally with the ratio of the total tempera¬ ture at the face of the engine divided by the maximum total design temperature at the face of the engine at sea level static design conditions. Rotor speed for the con¬ ventional engine at values of Θ ₂ above 1 is substantially constant, and at values of Θ₂ below 1 varies with the same ratio as does the turbine temperature.

OMPI FIGURE 5 depicts three graphs of the engine corrected airflow of the conventional turbojet engine and the point design engine operated at an intermediate throttle set¬ ting at three different altitudes (sea level, 20,000 feet and 36,089 feet). The graphs show that with increasing Mach number, the engine corrected airflow of the conven¬ tional SLS engine necessarily drops once the turbine in¬ let temperature and actual compressor speeds are at their maximum allowable values. On the other hand, the engine corrected airflow for the point design engine of the present invention remains substantially constant for all flight Mach numbers up to the point design flight Mach number of 2.0. It will be noted that the graphs for the point design engine show a slight increase in engine corrected airflow with increases in Mach number at all three altitudes. This slight increase is caused by the computer simulation technique employed to generate the curve shown in^" FIGURE 5. Under actual operating condi¬ tions, however, the engine corrected airflow would be maintained at or near the predetermined constant value and would not vary nearly as much as shown in the graphs. The engine corrected airflow is maintained at a constant value in the point design at flight Mach numbers below 2.0 by reducing the turbine inlet temperature along with a decreasing compressor inlet temperature to prevent corrected rotor overspeed. The reduction in turbine in¬ let temperature is scheduled to hold the maximum engine corrected airflow at a constant compressor match point for all flight conditions.

The turbine inlet temperature schedule required to main¬ tain the constant corrected airflow and to prevent

OMPI

. wipo corrected rotor overspeed for the point design engine of the present invention as illustrated in the graphs of FIGURE 6. Referring to FIGURE 6, the dot-dash line graphs of turbine inlet temperature correspond to the altitudes of sea level, 20,000 feet and 36,089 feet.

Since the conventional turbojet is designed for maximum performance at an intermediate throttle setting at sea level static, it is designed for operation at its maxi¬ mum turbine inlet temperature at sea level static con- ditions. Since the turbine inlet temperature cannot be allowed to exceed that maximum without damage to the en¬ gine, the turbine inlet temperature remains constant at the maximum as the flight Mach number increases, causing the engine porrected airflow to drop with increasing flight Mach number. The decrease in engine corrected airflow affects the engine performance, since the net thrust output of the engine falls off in proportion to the decrease in airflow. When the conventional turbojet engine is operating at 20,000 feet, the engine corrected airflow remains substantially constant until the turbine inlet temperature reaches its maximum allowable value at a flight Mach number of about 1.0. With increasing flight Mach numbers, the turbine inlet temperature thereafter remains constant and is accompanied by a de- crease in engine corrected airflow as shown in FIGURE 5. Likewise, when the conventional turbojet engine is oper¬ ating at 36,089 feet and above, the turbine inlet temp¬ erature is below the maximum allowable value at flight Mach numbers below about 1.2, but reaches the maximum at a flight Mach number of1.2 At flight Mach numbers above 1.2, engine corrected airflow again begins to de¬ crease with increasing flight Mach numbers as shown in FIGURE 5._ Referring to FIGURE 7, the uninstalled intermediate thrust (nonafterburning) versus flight Mach number for the conventional turbojet engine and the point design engine of the present invention is graphically illus- tra ed for various altitudes. At all altitudes, the un¬ installed net thrust of the point design engine of the present invention is lower than that of the conventional turbojet engine through the initial portion of the flight regime. However, at the upper end of the flight regime, that is, at the top speeds for a given altitude, the uninstalled net thrust of the point design engine exceeds that of the conventional turbojet engine. The intermediate thrust of the point design engine is greater than that of the conventional engine at the upper end of the flight regime for a given altitude because of the higher corrected airflow capability of the point design engine, as discussed in conjunction with FIGURE 5. At lower flight Mach numbers, the net thrust of the point design engine of the present invention falls below that of the conventional turbojet engine as the turbine inlet temperature of the point design engine is reduced. As discussed in conjunction with FIGURE 6, the reduction in turbine inlet temperature causes the lapse in intermedi¬ ate thrust with decreasing flight Mach numbers for the point design- engine relative to the conventional design.

The lapse in thrust at lower flight Mach numbers in the point design engine of the present invention is accom¬ panied by a decrease in uninstalled specific fuel con¬ sumption. Referring to FIGURE 8, the dashed lines graph- ically illustrate the specific fuel consumption of the

-ρUREΛT

OMPI conventional turbojet engine at altitudes of 20,000 feet and 36, 089 feet and above, while the solid lines graph¬ ically illustrate the specific fuel consumption of he point design turbojet engine with increasing Mach number. At 20,000 feet, the spread between the point design en¬ gine and the conventional turbojet engine is on the order of 0.2 pounds per pound per hour (lb/lb -hr) indicating that, although there is a lapse in performance by the point design engine, there is an advantageous result in that the fuel consumption of the point design engine is not as high as the conventional engine. Likewise, at 36,089 feet and above, the difference in specific fuel consumption between the point design turbojet ^"engine and the conventional turbojet engine is between 0.1 and 0.2 Ib/lb-hr, again indicating lower fuel consumption by the point design engine. The improvement in the specific fuel consumption of the point design engine relative to the conventional engine, especially at higher Mach num¬ bers, is due to the improved propulsion efficiency of the point design engine.

Since the intermediate thrust lapse of the point design engine of the present invention may be too severe for acceptable levels of airplane acceleration, especially through the transonic flight regime, it may be desirable to augment the intermediate performance of the engine by employing an afterburner, such as described in con¬ junction with FIGURE 4. For purposes of comparing the augmented performance of the point design engine and a conventional SLS turbojet engine, the maximum augmented thrust for an afterburner in each of the engines that operates at a constant exit temperature of 3800° Rankine at all flight Mach numbers was estimated at several alti¬ tudes over a given flight speed range. As illustrated in FIGURE 9, the uninstalled augmented thrust (in 1,000's of pounds) for the point desig engine and the conven¬ tional engine is separately plotted versus increasing Mach number for altitudes of sea level, 20,000 feet and 36,089 feet and above. The solid line graphs represent the augmented thrust of the point design engine while the dashed lines represent the augmented thrust of the conventional turbojet engine. At high Mach numbers, the point design engine has greater augmented thrust, due to the increased airflow handling capabilities. However, at lower flight Mach numbers, the augmented thrust of the conventional engine is superior to that of the point design engine. Although now shown by graphic illustration, it must be realized that the uninstalled specific fuel consumption at an augmented thrust setting is significantly higher for the point design turbojet engine when compared to the conventional turbojet. How¬ ever, at the highest Mach numbers, the specific fuel consumption of both engines operating at an augmented throttle setting converge to provide almost identical specific fuel consumptions..

The installed performance of the point design engine has also been estimated and compared with the conventional engine. The installed thrust at intermediate throttle settings and at augmented throttle settings is very similar to that calculated for uninstalled performance as graphically illustrated in FIGURE 9. Furthermore, the installed specific fuel consumptions are very similar to the uninstalled specific fuel consumptions when a fixed geometry inlet configuration is employed.

From the foregoing data, it can be readily ascertained that the point design turbojet engine offers large in¬ creases in intermediate thrust over the conventional turbojet engine at flight Mach numbers on the order of 2.0. The high intermediate thrust levels provide the capability to achieve and sustain supersonic cruise speeds at intermediate thrust levels rather than relying on continuous augmented thrust levels as required with conventional turbojet engines. The tradeoff that must be accepted with the point design engine is the inter- mediate thrust lap in the transonic regime relative to the conventional engine. As discussed, however, the intermediate thrust lapse can be largely overcome by augmenting the thrust to achieve a high acceleration rate through the transonic flight regime. The augmented thrust for the point design engine of the present in¬ vention offers comparable thrust levels for transonic acceleration to that of the conventional engine while still retaining a high level of intermediate thrust at supersonic cruise. Thus the overall benefits of the point design engine outweigh the disadvantages of the conventional turbojet engine for extended supersonic cruise conditions, primarily because the specific fuel consumption is lower at cruise speeds. Perhaps the most important advantage to the point design engine of the present invention is that acceptable thrust levels can be achieved in all flight regimes from sea level to 60,000 feet or more and from flight Mach numbers up to on the order of 2.0 with fixed inlet geometry.

Referring now to FIGURE 10, a second engine point de- signed for maximum efficiency at intermediate thrust settings at a flight Mach number of 2.7 is illustrated. The engine, similar in most respects to that previously described, is a single spool turbojet having single com¬ pressor stage and one turbine stage having variable angle rotor blades. The engine is designed, for example, to have a compressor pressure ratio on the order of 8 and a turbine inlet temperature on the order of 2,800°F. The engine employs an afterburner with a variable geom¬ etry exhaust nozzle. To achieve maximum thrust for acceleration, it is preferred that the engine be opera¬ ted with augmented thrust at takeoff to minimize the field length required for takeoff and that the engine be operated with augmented thrust at transonic speeds to achieve maximum transonic acceleration. However, once the flight Mach number approaches the cruise speed de¬ sign point of 2.7, the augmented thrust can be eliminated and only intermediate thrust employed to achieve and maintain a cruise speed of Mach 2.7. Still referring to FIGURE 10, the Mach 2.7 engine 130 is illustrated in conjunction with a fuel flow control system for main¬ taining the maximum corrected airflow through the engine constant at flight Mach numbers below 2.7. The engine 130 includes a housing 132, inlet face 134 and a vari¬ able geometry nozzle 136. The engine has a single spool including a drive shaft 138 mounted for rotation

on bearings 140 that directly couple the five compressor rotors 142, 144, 146, 148 and 150 with the single turbine rotor 152. Each of the compressor rotors carries a set of compressor blades 142a through 150a. Stator blades 154, 156, 158 and 160 are associated with each of the rotors in a conventional manner. An annular combustor, generally designated 162, is located in the engine aft of the compressor. The combustor 162 houses the nozzles for spraying fuel into the combustor section. A single spray bar 164 is shown in FIGURE 10 for injecting fuel into the combustor 162. The turbine rotor 152 is situ¬ ated aft of the combustor 162 and includes variable angle rotor blades 152a, .which will be described in greater detail in conjunction with FIGURE 11. Station- ary turbine guide vanes 163 are situated forwardly of the rotor blades 152a. The combusted gases from the combustor 162 pass through the guide vanes 163 and ro¬ tate the turbine rotor 152 by reaction with the turbine blades 152a. After the combusted gases are expanded through the turbine, they pass into the afterburning section 166 of the engine. Fuel is supplied to the afterburning section via an afterburner spray bar 167 (only one of which is shown in the drawings) supplied with fuel via line 170 from a fuel source (not shown) . The fuel supply through line 170 is regulated in a con¬ ventional manner by conventional valving apparatus re¬ sponsive to pilot controlled throttles. When the after¬ burner is employed with the engine 130, it is preferred to also employ a variable geometry nozzle 136 that is normally oriented in a first position shown in dot-dash outline 136a for intermediate throttle settings to re¬ strict the nozzle throat. When the afterburner is in

iUR _Al

OMPI operation, the throat area of the nozzle is increased by moving the variable nozzle structure radially out¬ wardly to the position shown in solid outline. The purpose of the variable geometry nozzle is the same as that for the previously described engine, namely, to relieve back pressure on the turbine and thus to match the flow and pressure of the nozzle to the corresponding speed and temperature of the turbine at intermediate throttle settings and augmented throttle settings. The control and actuation of the variable geometry nozzle 136 are now shown, as any of a variety of conventional variable geometry nozzles and actuating mechanisms can be employed with the engine 130.

The total inlet temperature at the face of the engine and, similarly, the engine corrected airflow, are main¬ tained at a constant value for intermediate throttle settings for all flight Mach numbers below, the point de¬ sign Mach number of 2.7 by varying the fuel input to the engine to decrease the physical rotor speed as the flight Mach numbers decrease from the point design Mach number. In this embodiment of the invention, as illus¬ trated in FIGURE 10, the engine operating parameter that is correlated to and representative of the engine cor¬ rected airflow is the compressor pressure ratio. The compressor pressure ratio can be sensed by positioning a first pressure sensor 168 at the compressor inlet and a second pressure sensor 169 at the compressor outlet. Pressure signals provided by the pressure sensors 168 and 169 are transmitted via lines 172 and 174 to a com- parator and signal conditioner 176. The comparator 176

- UREA

QMPI_ compares the turbine inlet and outlet pressure signals provided through lines 172 and 174 and yields a differ¬ ential signal that is indicative of the pressure ratio across the compressor of the engine. The differential signal is internally conditioned in the comparator/signal conditioner 176 to yield an output signal representative of the actual engine corrected airflow.

The output signal indicative of engine corrected airflow is transmitted via line 178 to the second comparator 180 where it is then compared with a reference signal input that corresponds to the maximum point design engine cor¬ rected airflow. The second comparator 180 provides an output control signal representing the difference between the actual and reference engine corrected airflow signals. The output signal from the second comparator 180 is trans¬ mitted via line 186 to control apparatus for adjusting the turbine inlet temperature to equilibrate the actual engine corrected airflow with the desired or reference engine corrected airflow in a manner similar to that of the engine described in conjunction with FIGURE 4. For example, the output signal transmitted via line 186 can be employed to energize a servocontrol 188, which in turn actuates a fuel flow override control valve 190 to appropriately adjust the fuel flow into the combustor 162 of the engine 130 and thus appropriately adjust the turbine inlet temperature. Fuel is conventionally sup¬ plied to the engine 130 from a source (not shown) through fuel line 192. The fuel flow rate is initially control¬ led by a manually operated fuel flow control valve 194 before entering the override flow control valve 190. The manually operated fuel flow control valve 194 is actuated by manually controlled throttles in a manner similar to that described in conjunction with the engine of FIGURE 4. Thus at a manually set full intermediate thrust setting, the fuel flow control valve 194 is opened to its maximum extent. Fuel from the manual fuel flow control valve 194 is then variably restricted, dependent upon the compressor pressure ratio, by the override flow control valve 190 to meter the amount of fuel entering the combustor through the spray bar 164 and thereby main¬ tain the desired engine corrected airflow through the engine at all flight speeds.

Referring to FIGURE 11, in order to provide an engine that is point designed to provide maximum intermediate thrust at a cruise Mach number of 2.7 and that will pro¬ vide adequate thrust at intermediate throttle settings in the transonic flight regime, it is preferable to em¬ ploy variable turbine geometry. As• illustrated in FIG- URE 11, the turbine guide vanes 163 are located ahead of the turbine rotor blades 152a. The rotor blades 152a are fixed, while the guide vanes 163 are mounted on radially oriented blade shafts 152b, in turn mounted for rotation on the engine housing forwardly of the turbine rotor 152, so that the angle of the guide vanes 163 can be varied relative to the angle of the rotor blades 152a. As schematically shown, the guide vane position can be varied between the position shown in solid outline where the angle of the rotor blades is small relative to the guide vanes and the position shown in dotted outline where the guide vanes are oriented at a greater angle relative to the rotor blades 152. The vane shafts 163a are coupled by a suitable actuating mechanism to a servo¬ control 198. The vane angle is changed throughout the flight regime to achieve the desired turbine inlet temp- eratures and corrected airflow through the engine. At lower flight Mach numbers (at subsonic flight speeds) , the vanes -are positioned at the lesser angle (shown in dotted outline in FIGURE 11) . As the flight speed is increased, the vane angle is increased so that at super- sonic flight speeds approaching the Mach 17 design point the vanes will be positioned at the greater angle (shown in solid outline) . Although a variety of engine opera¬ ting parameters and flight conditions can be sensed and used to control the blade position, the simplest param- eter to measure and implement for control is the flight total temperature. The flight total temperature is sensed by a flight condition sensor 200. The flight con¬ dition sensor generates a signal that is indicative of the flight total temperature at a given flight Mach number. The signal is transmitted to a comparator 204 via a line 202. The comparator compares the sensed flight condition signal with a reference signal input to the comparator that is indicative of the desired vane angle, and generates a differential control signal that is transmitted via line 206 to selectively energize the servocontrol 198. The servocontrol motor actuates the vane shaft to change the vane angle to the desired posi¬ tion for the given flight condition sensed.

Claims

1. A supersonic airplane with a turbine engine designed to provide a predetermined amount of thrust at a given throttle setting when propelling said airplane at a pre¬ determined supersonic Mach number in the range of from about 2.0 to about 2.7, the improvement comprising:

control means for maintaining the corrected airflow in said engine substantially at a constant value at flight Mach numbers below said predetermined supersonic Mach number; and

wherein said turbine engine has a fixed geometry inlet for ducting air to said compressor.

2. The supersonic airplane of Claim 1, wherein said turbine engine includes an afterburner and a variable geometry exhaust nozzle for directing exhaust gas away from said engine.

3. The supersonic airplane of Claim 2, wherein said turbine includes fixed turbine blades and stator vanes, said improvement further comprising means for mounting said vanes for rotation about an axis oriented substan- tially radially relative to said drive shaft, said vanes being so constructed and mounted as to be movable be¬ tween a first position wherein said vanes are oriented at a first angle relative to said turbine blades and at least a second position wherein said vanes are oriented at a second angle relative to said turbine blades that is greater than said first angle, and wherein said im¬ provement further comprises means for rotating said blades from said first position to said second position when a predetermined flight condition of said aircraft exceeds a predetermined value.

4. A method for controlling a turbine engine point de¬ signed to provide a predetermined amount of thrust at a given throttle setting when propelling an aircraft at a predetermined supersonic Mach number, comprising the steps of: controlling the corrected airflow in said engine and maintaining the corrected airflow constant at flight Mach numbers below said predetermined super¬ sonic Mach number.

5. The method of Claim 4, wherein said turbine engine includes a compressor, a turbine, and a single drive shaft interconnecting said turbine and said compressor, and wherein the corrected airflow in said engine is maintained constant by reducing the rotational speed of said drive shaft as the flight Mach number is lowered below said predetermined supersonic Mach number.

6. A method for controlling a turbine engine point designed to provide a predetermined amount of thrust at a given throttle setting when propelling an airplane at a predetermined supersonic Mach number of about 2.0, said turbine engine including a compressor, a turbine, and a single drive shaft interconnecting said turbine and said compressor, said turbine having fixed turbine blades, comprising the steps of:

OMPI

^' sensing an engine operating parameter to provide a first signal representative of the corrected airflow in said turbine engine,

comparing said first signal with a second signal repre- sentative of a predetermined maxim-urn value of said en¬ gine operating parameter,

providing a control signal representative of the differ¬ ence between said first and second signals, and

adjusting the engine operating condition of said engine in response to said control signal to maintain the cor¬ rected airflow through said engine constant at flight Mach numbers below.2.0.

7. The method of Claim 6, wherein said engine operating parameter is a pressure ratio across a predetermined portion of said engine.

8. In a method for controlling a turbine engine point designed to provide a predetermined amount of thrust at a given throttle setting when propelling an airplane at a predetermined supersonic Mach number of about 2.7, said turbine engine including a compressor, a turbine, and a single drive shaft interconnecting said turbine and said compressor, said turbine having fixed turbine blades, including the steps of:

sensing an engine operating parameter to provide a first signal representative of the corrected airflow in said turbine engine,

-BU EA

OMPI comparing said first signal with a second signal rep¬ resentative of a predetermined maximum value of said engine operating parameter,

providing a control signal representative of the differ- ence between said first and second signals,

adjusting the engine operating condition of said engine in response to said control signal to maintain said en¬ gine operating parameter constant at flight Mach numbers below 2.7,

an improvement comprising the steps of: sensing the flight total temperature; and increasing the angle of the turbine stator vanes relative to the turbine blade at flight Mach numbers below a predetermined flight Mach number less than 2.7.

9. The method of Claim 8, wherein said engine oper¬ ating parameter is a pressure ratio across a predeter- mined portion of said engine.

10. An airplane having a wing, a vertical stabilizer, a canard, and a turbine engine mounted in the fuselage of said airplane, said engine being point designed to provide maximum thrust at a given throttle setting when said airplane is being propelled by said engine at a predetermined supersonic flight Mach number, said engine control means for maintaining the corrected airflow through said engine at a substantially constant value

-gU EA T

OMPI at flight Mach numbers below said predetermined super¬ sonic flight Mach number, said airplane including an inlet positioned adjacent the rearward portion of said canard for ducting air to said turbine engine, and a nozzle located in the aft portion of said fuselage for directing exhaust gases from said engine rearwardly from said engine and said fuselage. -

11. The airplane of Claim 10, wherein said canard is located on said fuselage above and forwardly from said inlet and is positioned relative to said inlet so as to provide precompression of the air flowing under said canard and into said inlet while said airplane is in flight.

12. The airplane of Claim 11, wherein said predeter¬ mined supersonic flight Mach number is in the range of from about 2.0 to about 2.7.

13. The airplane of Claim 12, wherein said inlet has fixed geometry and wherein said exhaust nozzle has vari¬ able geometry, said predetermined supersonic flight . number being about 2.7.

OMPI