JET ENGINE TESTING APPARATUS

Abstract: Apparatus for detecting jet engine failures during aircraft flight. Low and high speed rotor rotation velocities of one engine are compared with low and high rotor rotation velocities of other engines and the difference between them utilised to determine engine failure. Equivalent low speed rotor velocity of an engine is computed from the high speed rotor velocity of that signal. This computed value is compared with measured low speed rotor velocity of the same and of different engines. Differences between the computed values and the actual values are compared to a computed threshold, which allows for normal differences between engines, to determine the operational state of the engine. Additional engine testing is performed by comparing low speed rotor rotational velocity to a computed windmilling value for each engine.

Description: This present invention relates generally to jet engine testing, and more specifically to failure detection of jet engines on aircraft during all phases of flight. Many commercial transport aircraft, general aviation aircraft, and military aircraft, are powered by turbojet or turbofan engines. Computers on board the aircraft are utilised for the determination of engine failures from monitored engine parameters. Computer command speeds and thrust settings for safely flying the aircraft, after a failure has been detected, are displayed to the human pilot or provided to an automatic flight control system. Prior art jet engine failure detection systems monitor Engine Pressure Ratio (EPR) and Exhaust Gas Temperature (EGT). EPR is derived from pressure sensors mounted in the air inlet and the exhaust nozzle of an engine. The ratio of these two pressures is a direct measure of the thrust of the engine, whereas EGT is derived from a temperature sensor mounted in the exhaust nozzle of an engine. Systems utilising these parameters for jet engine failure detection have numerous short-comings. EPR cannot be used during the descent of an aircraft, because the EPR during this phase of flight is substantially equal to the EPR of the failed engine. Additionally, the EPR while an aircraft is cruising is small, making the determination of EPR level at which an engine may be considered to have failed, during this phase of flight, extremely difficult. Furthermore, the air quality and temperature in the operating regions may cause pressure sensors positioned in the air inlet to become clogged with ice or debris, thereby producing erroneous measurements. Measuring EGT also does not provide reliable jet engine failure indications. For example, the EGT of an engine on fire remains near that of an normally operating engine for an appreciable time period after the start of the fire. Further, since residual heat is dissipated into the atomosphere, the EGT of an engine that has failed for reasons other than fire, is an exponential decay exhibiting a long time constant. Consequently, a significant time lapse may occur before an alarm is given. Another deficiency of the EGT method results from the fact that different models of the same engine may exhibit appreciably different EGTs at identical thrust settings, thereby creating the possibility of a false failure indication. The limitations inherent in the use of EPR and EGT engine parameters cause failed engine detection to be flight phase dependent, i.e. a different set of criteria must be used for takeoff, cruise, descent etc. Engine failure detection systems utilising these measurements exhibit high probabilities of erroneous failure detection when the aircraft transitions from one phase of flight to another, for example changing from climb to cruise. Most modern day turbojet or turbofan engines utilise two rotor shafts, the first, a low speed rotor, rotating at N₁ revolutions per minute and the second, a high speed rotor, rotating at N₂ revolutions per minute. A compressor stage, for compressing incoming atmospheric air, is positioned on each rotor shaft near the air inlet end, while a turbine is mounted on the rotor shaft at the end opposite the compressor. As the exhaust gas exits the engine, it passes through the turbine blades causing the rotor to rotate, thereby driving the compressor section. Generally, rotor angular velocity is expressed as a percentage of the maximum allowable rotational velocity. The determination of the high and low speed rotor angular velocities allows engine failure detection to be completely independent of the flight phase, thus permitting the use of constant engine failure criteria for all flight phases. The rotation velocities N₁ and N₂ are inter-related, one being determinable from the other, for any thrust setting, by a well defined function. This relationship may be utilised to compare the computed and measured shaft rotation speeds for an engine. Additionally, the shaft rotation speed of one engine may be compared to the shaft rotation speed of all other engines mounted on the aircraft. Other turbojet, turbofan engine characteristics may be calculated or measured for the determination of engine failure. In turbojet, turbofan engines rotors of a failed engine, that are not jammed, will rotate due to the forward motion of the aircraft. This rotation, known as windmilling, has an angular velocity which is a linear function of the aircraft Mach number, thereby providing an additional parameter for detection of a failed engine. US-A-4,215,412 discloses an engine monitoring system which uses current values of various engine operating parameters to predict the current values of a set of engine performance parameters. Reference is made to the possible use of a dependent parameter in the form of the rotational speed of a low pressure compressor by determining that parameter from an independent parameter using linear regression analysis. More particularly, there is disclosed apparatus for testing jet engines mounted on an aircraft, the jet engines being of the type having a high speed rotor and a low speed rotor, the apparatus including difference means for obtaining signals representative of the differences between signals applied thereto; means for providing first signals representative of rotation rates of the low speed rotor in the jet engines; means for providing second signals representative of rotation rates of the high speed rotor in the jet engines; means for providing threshold signals; and engine failure indicator means. The present invention is defined in the appended claims and jet engine testing apparatus in accordance with the invention includes sensing devices that provide electrical signals representative of the rotation rates of the high and low speed rotors in a turbofan or turbojet engine. The high and low speed rotor representative signals for each aircraft are respectively subtracted from the high and low speed rotor representative signals of a selected one engine of all the engines on the aircraft to obtain high and low speed rotor difference signals. The level of these difference signals are respectively compared to high and low speed rotor threshold signals derived from signals representative of the maximum high and low speed rotor rotation rate of all the jet engines on the aircraft. Identification of an engine that has failed is determined from the comparison of the difference signal level with the threshold signal and the polarity of the difference signal. Each engine is also checked by deriving a signal from the representative signal of the high speed rotor rotation rate that is equivalent to the representative signal of the low speed rotor rotation rate of a normally operating engine. This equivalent signal is subtracted from the signal representative of the low speed rotor rotation rate of the reference engine to derive a difference signal that is compared to the low speed rotor rotation rate threshold signal. The level and polarity of each of these difference signals is determined and utilised to establish and identify failed engines. Further testing for engine failure is performed by comparing the signals representative of the low speed rotor rotation rate to a signal, derived from the forward velocity of the aircraft, that is representative of the rate at which the low speed rotor will rotate due to the flow of air through the engine at the forward velocity of the aircraft. An engine failure indication is given when this comparison shows that the calculated windmilling rotation rate exceeds or equals the sensed rotation rate of the low speed rotor. Jet engine testing apparatus in accordance with the present invention will now be described in greater detail, by way of example, with reference to the accompanying drawings in which:- Figure 1 is a simplified cross-sectional view of a typical turbojet engine useful for explaining the general operation of the engine, Figure 2 is a graph illustrating the relationship between turbojet engine windmilling revolution rate and the Mach number of the aircraft, Figure 3 is a graph illustrating the relationship between the low and high speed rotor rotation rates for a properly operating turbojet engine, Figure 4 is a block diagram of the circuitry utilised for determining the integrity of the rotation rate sensors, Figures 5a and 5b are block diagrams of the circuitry utilised for deriving threshold signals, Figure 6 is a block diagram of circuitry utilised for deriving signals representative of the rotation rate of the low speed rotor, of a normally operating jet engine, corresponding to a sensed rotation rate of the high speed rotor, Figure 7 is a block diagram of circuitry utilised for determining the windmilling rotation rate of a jet engine from measured forward aircraft velocity, Figures 8a and 8b are block diagrams of circuitry utilised for comparing rotation rate of corresponding rotors in two engines to test for engine failure, and Figure 9 is a block diagram of circuitry utilised for the comparison of measured low speed rotor rotation rates to computed low speed rotor rotation rates and to computed windmilling rotation rates. The engine failure detection scheme of the present invention utilises the measurements of the low and high speed rotor shaft rotational velocities of a turbojet or turbofan engine. While the present invention may be used on aircraft having any number of engines, for the sake of clarity of explanation, the following discussion will assume an aircraft having just two engines. Figure 1 is a simplified longitudinal cross-sectional view of a typical jet engine 10. Under normal conditions, atmospheric air enters the engine through an inlet 11 to be compressed by a low speed compressor 12 which is physically connected to a low speed rotor shaft 13. Compressed air from the low speed compressor 12 is passed, for further compression, to a high speed compressor 14, which is physically connected to a high speed rotor shaft 15. Highly compressed air from the high speed compressor 14 is then passed to a combustion chamber 16, wherein the air is mixed with fuel and ignited. Combustion gas products pass through a high speed turbine 17, physically attached to the high speed rotor shaft 15, causing rotation of the high speed turbine 17, the high speed rotor shaft 15, and the high speed compressor 14. These gases also pass through a low speed turbine 18 causing rotation of the low speed turbine, the low speed rotor shaft 13, and the low speed compressor 12 prior to exiting the engine through an exhaust nozzle 19 to produce the engine thrust. It should be apparent from this illustration that the high and low speed stages are mechanically decoupled, the action of one affecting the action of the other only by the flow of air through the engine. The rotational velocity of the low and high speed rotor shafts 13 and 15, N₁ and N₂ respectively, may be measured by a tachometer or any other suitable means and supplied to an onboard computer in units of percentage of the maximum allowable rotational velocity. These measured engine velocities may be utilised in the computer for the determination of an engine failure. A turbojet or turbofan engine may fail in any of a number of ways, the most common being failure of the fuel/air mixture to ignite. When this occurs combustion gases have relatively low energy and cannot maintain sufficient rotational velocities for the turbines 17 and 18. Neither rotor, however, ceases to rotate since the air forced through the engine by the forward velocity of the aircraft causes both to windmill at a rotational speed that is proportional to the Mach number of the aircraft forward velocity. The relationship between the rotational velocity of the low speed rotor due to windmilling and the forward velocity Mach number is shown in Figure 2, wherein the vertical axis 21 is the low speed rotor shaft rotational velocity expressed in percent of the maximum allowable rotational velocity, while the horizontal axis 22 is the aircraft Mach number. Straight line 23 represents the relationship between the two parameters, expressed as: (1) N₁ = k₁M where: N₁ = the rotational velocity of the low speed rotor shaft due to windmilling expressed in percent of maximum allowable velocity. k₁ = a proportionality constant. M = the aircraft Mach number. Though the high and low speed rotors are not mechanically coupled, the action of the combustion gases passing through the high speed turbine 17 and low speed turbine 18 cause the two rotors to rotate at angular velocities that are approximately linearly related. This approximate linear relationship, shown in Figure 3 as line 24, may be expressed in terms of the percentages of the maximum allowable high and low speed rotation rates by the equation: (2) N₁ = K₂ N₂ - C where K₂ and C are appropriate constants. This relationship and the windmilling rotation rate dependence on the forward velocity Mach number of the aircraft provide criteria for the determination of the failed engine in the present invention. Properly operating measurement equipment is required to provide reliable engine fault detection. A block diagram of an apparatus for testing the rotation sensors prior to take-off is presented in Figure 4. A signal representative of the N₁ measurement obtained from a low speed rotation sensor 30 under test, which may be a tachometer with a frequency-to-voltage converter, is supplied via a lead 32 to a level detector 33. The level detector 33 provides a logic 1 via a lead 34 whenever the N₁ representative signal exhibits a value below a threshold signal that is representative of a selected rotational rate, as for example 15% of maximum allowable rotational velocity, and a logic zero otherwise. In a similar fashion, a signal representative of the N₂ measurement from a high speed rotation sensor 31 under test is coupled via a lead 35 to level detector 36, wherefrom a logic 1 is coupled via a lead 37 whenever the N₂ representative signal indicates a value below a selected threshold, and logic zero otherwise. The logic signals on the leads 34 and 37 are coupled to a conventional exclusive OR gate 38, wherefrom a logic 1 is coupled to AND gates 42 and 43 via leads 44 and 45 when one, and only one, of the logic signals on leads 34 and 37 is a logic 1, otherwise the signals coupled to the leads 44 and 45 are logic zeros. Second input terminals of the AND gates 42 and 43 are respectively coupled to receive the logic signals from level detectors 33 and 36. Since each AND gate 42 and 43 provides a low level logic signal when a low level signal is coupled to either one of the two input terminals, and a high level signal when high level logic signals are coupled to both input terminals, a logic 1 is coupled to a lead 46 whenever the rotational velocity measurement for the low speed rotor indicates a low speed rotor velocity below the said threshold. In a like manner a logic 1 is coupled to the lead 47 whenever the rotational velocity measurement for the high speed rotor indicates a high speed rotor velocity below the threshold. Those skilled in the art recognise that a valid measurement is indicated when both the high and low speed rotors fail simultaneously. This ambiguity is readily resolved for it would not be possible for the aircraft to take off under these conditions. Further, the probability of such an occurrence is extremely small. Similar logic may be employed to determine rotational velocity indications that are higher than the actual rotational speed of the rotors. This logic utilises the threshold signal amplitude above the maximum allowable rotation velocity, as for example 15%, to provide a logic 1 output from the level detectors when the rotational velocities exceeding this threshold are indicated. an indication of a valid measurement may occur when both sensors are providing erroneously high signal amplitudes. The probabilities, however, of both sensors simultaneously providing speed indications greater than 15% above the actual rotor speed is an insignificantly low value. Concurrently with the action described above, a signal proportional to the aircraft Mach number is coupled from an air data computer 51 to a third level detector 52, the latter operating to provide a logic 1 output signal when the Mach number of the aircraft velocity exceeds a predetermined value, as for example Mach 0.12, a speed readily achieved during taxi, and a logic zero otherwise. This logic signal is coupled to latches 53 and 54 via lines 55 and 56, respectively. The latches 53 and 54 latch the signals respectively coupled thereto from the lines 46 and 47 and maintain these logic signals for the duration of a high level logic signal coupled from the level detector 52. Once high level logic signals simultaneously appear at the input terminals of the latches 53 and 54, these latches maintain a high level signal at the respective output terminals 57 and 58, even though a low level logic signal is subsequently coupled from the sensor test circuitry. Simultaneously high level logic signals from the AND gates 42 and 43 are an indication that one of the sensors has failed. When the aircraft begins to accelerate for takeoff, its speed increases until it exceeeds the Mach number threshold level set for the level detector 52. At this time a high level logic signal is coupled to the latches 53 and 54 forcing each to maintain logic signals coupled via the leads 46 and 47 respectively. It is reasonable to assume that the aircraft would not accelerate for takeoff with an engine that has failed. Thus it may be reasoned that the latches were set properly. It is also reasonable to assume that not both N₁ and N₂ measuring devices have failed Consequently, high level logic signals simultaneously coupled from the AND gates 42 and 43 signify that the N₁ measuring logic has failed or that N₂ measuring logic has failed. In this manner, the status of the two measuring devices is determined before the aircraft has left the ground. Rotor rotation threshold levels may be determined for the engines on the aircraft. Differences in throttle lever-handle-to-engine fuel control valve rigging and differences in fuel control valves from engine to engine cause significant variations in rotor rotational velocities with thrust increase commands. Often a given engine would achieve a desired thrust setting significantly before another engine of the same type. The threshold levels are functions of the maximum measured RPM and the idle RPM of the engines and are set to account for the variations noted above. A block diagram of, and apparatus for, determining the threshold values for the low speed rotor is shown in Figure 5a. The signals representative of N₁ from the first and second engine measuring devices 70, 71 are coupled to a maximum level detector 72, wherein the greater of the two N₁ representative signal levels is selected and coupled to an amplifier 74, having a gain that is representative of a desired scale factor. Signals at the output terminals of the amplifier 74 are subtracted from the output signals of a constant signal generator 75 in a difference circuit 76. The amplifier 74 gain and the level of the constant signal are chosen to provide threshold signals at the output of the difference circuit 76 which are linearly proportional to the maximum measured N₁ signal and exhibits a zero crossing at a desired fraction of the rotation rate of the maximum measured N₁ signal. To establish a rotation rate range about the desired rotation rate, this difference signal is coupled to a limiter 79, wherefrom signals are coupled to an output terminal 80 that are between two desired limits, as for example between signals representing 25% and 15% of the rated maximum rotation rate for the low speed rotor with a zero crossing at 20% of the rated maximum rotation rate. The signals coupled to the output terminal 80 are thresholds used in a subsequent process, to be described, for comparing the N₁ measurements of other engines. Threshold values for the N₂ measurements signals are established in a similar manner. Referring to Figure 5b, N₂ representative signals for the high speed rotors of the engines are coupled from N₂ measuring devices 81, 82 to a maximum level detector 83. The signal representative of the maximum N₂ rotational velocity is coupled to the difference terminal of a difference circuit 84 via an amplifier 85, while the constant level signal is coupled to a positive terminal at the difference circuit from a constant signal generator 86. This constant signal may be representative of twice the rated maximum rotational velocity of the high speed rotors. The constant and the gain of the amplifier 85 are chosen to provide threshold signals that are linearly proportional to the maximum measured rotational speed and a zero crossing at a desired fraction of the rated maximum speed. To establish a rotation rate range for the high speed rotors, the linearly varying signals are coupled to a limiter 89 which is set to constrain the signal levels coupled to the terminal 90 to be between two determined limits, as for example 40% and 20% of the maximum rate rotational speed of the high speed rotors with a zero crossing at 30% of maximum. Hence the magnitude of the signal at the terminal 90 represents a linear function constrained to values between the upper and lower values of the limiter 89 and is used in the determination of an engine failure based on N₂, a process to be discussed subsequently. A block diagram of a circuit that may be employed for deriving an average rotation rate of the low speed rotor, in a normally operating engine, from the measured rate of the high speed rotor is shown in Figure 6. The signal representative of the rotational velocity N₂ of the high speed rotor is coupled from a rotational velocity measuring device 91 to an amplifier 92 having a gain consistent with the coefficient of N₂ in equation 2. The output of the amplifier 92, which is equivalent to k₂ N₂, is coupled to a conventional summation network 93. A constant value signal representing the constant term in equation 2 is also coupled from a constant signal generator 94 to the summation network 93. The sum signal is coupled to a limiter 95 configured to provide output signals between levels corresponding to zero and 100% of the maximum nominal rotation velocity of the low speed rotor. Thus the output signal of the limiter 95 is representative of a nominal velocity of the low speed rotor corresponding to the measured velocity of the high speed rotor. The limiter 95 output signal is utilised in engine failure determination circuitry to be described subsequently. A signal representative of the speed at which the low speed rotor would windmill is established in a similar manner. Referring now to Figure 7, a signal from the air data computer 101 representative of the aircraft's forward velocity is coupled to an amplifier 102. The output signal from amplifier 102 representative of the term k₁ M in equation 1 is coupled to a conventional summation network 103. A signal at a constant amplitude representative of a selected low speed rotor rotation velocity, as for example 3% of the nominal maximum rotation velocity, is coupled to summation network 103 from constant level signal generator 104. This added constant value increases the windmilling representative signal from the amplifier 102 to allow for windmilling variations between engines. Output signals from the summation network 103 are coupled to a limiter 105 wherein the signal is limited to levels between 25% and 100% of the maximum windmilling rotation rate for an average engine. The output signals from the limiter 105 thus represent an engine's windmilling rotation rate, within the specified limits, at the forward velocity of the aircraft. The signals derived, as described above, are utilised to detect engine failures during flight and to identify which engine has failed. Referring to Figure 8a, the signals representative of the low speed rotor rotational velocity N₁ of first and second engines are coupled to a differencing circuit 106, wherefrom the difference of these signals, indicative of the difference between low speed rotor rotation velocities of the two engines, is coupled to the wiper arm terminal 107 of a switch 108. The wiper arm 109 of the switch 108 will be in the position shown when proper operation of the low speed rotor rotation sensors, for each engine, has been verified by the process previously described. If either or both of the N₁ measurements has been determined to be invalid, the wiper arm 109 is moved to make contact with a terminal 114, thus disabling the fault identifying circuit and preventing erroneous fault indications. When both N₁ measurements are valid the difference signal is coupled to a level detector 115 and a polarity detector 116. The detector 115 compares the absolute value of the difference signal to the absolute value of the signal, appearing at the terminal 80 Figure 5a, representative of a percentage of the maximum low speed rotor rotation velocity. If the difference signal coupled from the differencing circuit 106 exceeds the level of the signal at the terminal 80, the detector 115 couples a logic 1 signal to AND gates 117 and 118, otherwise a logic zero is coupled to the AND gates 117 and 118. The polarity detector 116 couples a logic 1 signal to the AND gate 117 and to an inverting terminal 119 of the AND gate 118 when the low speed rotor rotation velocity of the first engine is greater than low speed rotor rotation velocity of the second engine, and logic zero to the AND gate 117 and inverting terminal 119 when the low speed rotor rotation velocity of the second engine exceeds that of the first engine. Thus the gate 117 provides a high level logic signal to an engine failure indicator 120 when the first engine has failed, while the AND gate 118 provides a high level logic signal to the engine failure indicator 120 when the second engine has failed. The low speed rotor rotation velocity of the first or second engine may be compared in the same manner to the low speed rotor rotation velocity of other engines on the aircraft to determine which if any of the engines has failed. In a similar fashion the high speed rotor rotation velocity N₂ of the engines on the aircraft may be compared to provide additional information for the determination of which, if any, engine on the aircraft has failed. As shown in Figure 8b signals representative of the high speed rotation velocities are coupled to a difference circuit 121, the output signal of which is coupled to the wiper arm terminal 122 of a switch 123. A wiper arm 122a will be in the position shown when both N₂ measurements have been determined to be valid, by the process previously described, otherwise the wiper arm 122a will be in contact with terminal 124 to disable the fault indicating circuits. A level detector 125 receives the signal representative of the maximum rotation velocity of the high speed rotors from the terminal 90, Figure 5b. The level detector 125, polarity detector 126 and AND gates 127, 128 cooperate in the manner above described for the level detector 117, polarity detector 116 and AND gates 117, 118 to provide indications of engine failure. Figure 9 is a block diagram of a network for comparing the measured N₁ value of an engine (engine 1) with the value of N₁ computed from the measured N₂ for that engine, with the computed N₁ of another engine (engine 2), and with the calculated windmilling rotation rate for that engine (engine 1). Signals representative of the measured N₁ and N₁ as calculated from N₂ for an engine are coupled to a differencing network 131. The signal output from the differencing network 131 is proportional to the difference between the actual N₁ measurement and the N₁ computed from the measured N₂ rate. This difference signal is coupled to a switch 132, which is closed when the N₁ and N₂ measurements are determined to be valid, coupling the difference network 131 output signal to a level detector 133. If either measurement is determined to be invalid, the switch 132 is open, decoupling the level detector 133 from the differencing network 131. The level detector 133 compares the difference signal to the signal representative of the maximum measured low speed rotor rotation rate coupled from the terminal 80, Figure 5a. If the level of the difference signal exceeds the level of the maximum rate representative signal, the level detector 133 provides a logic 1, signifying a failure, via a lead 134 to the engine failure indicator 120 for further action. When the maximum rate representative signal exceeds the difference signal the logic 0, signifying proper engine operation, is coupled via a lead 134 to the computer. The signal representative of the measured N₁ value of engine 1 and the signal representative of the N₁ value of engine 2 computed from N₂ of engine 2 are coupled to a differencing network 141 and the difference signal therefrom is coupled to a switch 142, which operates in the same manner as the switch 132. The difference signal output from the network 141 is proportional to the difference of the measured N₁ for engine 1 and N₁ of engine 2 computed from N₂ of engine 2. This difference signal is coupled to the level detector 135 and polarity detector 136 which provide output signals, as previously described, with respect to level detector 115 and polarity detector 116 of Figure 8a. The output signals from level detector 115 and polarity detector 116 are coupled to the AND gates 137 and 138, which operate in the manner described with respect to the AND gates 117 and 118, to provide indications of engine 2 failure. As stated previously, when an engine fails the low speed rotor, due to windmilling, continues to rotate at a rate determined by the forward speed of the aircraft This windmilling rotation rate, computed as described above, may be utilised as another indicium of engine failure. Referring again to Figure 9, the calculated windmilling rate signal is coupled from the limiter 105 (Figure 7) to the positive terminal of the difference amplifier 151, while the signal representative of the low speed rotor rotation rate is coupled to the negative terminal. Thus the output signal of the difference circuit 151 is proportional to the difference between the actual measured N₁ and the computed windmilling value. This signal is coupled via a switch 152 to a polarity detector 153. The switch 152 operates in the same manner as the switch 132. The polarity detector 153 provides a high level signal at output terminals thereof, signifying an engine failure, when the measured N₁ is equal to, or less than, the computed windmilling rotation rate, otherwise the polarity detector 153 provides a low level logic signal, signifying normal engine operation. It should be appreciated by those skilled in the art that the present invention provides an improved method for timely and reliable detection of a failed jet engine. This is accomplished by determining allowable differences in rotor rotational velocities as a function of the maximum rotational velocity of the engines being tested to provide for normal differences in engine performance due to throttle handle rigging errors, differences in fuel control valves, etc.; comparing the difference in the low speed rotor rotational velocities of any number of engines against the computed allowable difference; comparing the high speed rotor rotational velocities of any number of engines against the computed allowable differences; comparing the difference between the low speed rotor rotational velocity and a computed low speed rotor rotation velocity to the computed allowable difference; and comparing low speed rotor rotation velocity to a computed windmilling value. The plurality of testing performed by the invention allows a single measuring device failure on each engine to be tolerated without significantly effecting the ability to detect an engine failure.

Claims: Apparatus for testing jet engines mounted on an aircraft, the jet engines being of the type having a high speed rotor and a low speed rotor, the apparatus including difference means (131) for obtaining signals representative of the differences between signals applied thereto; means for providing first representative signals representative of rotation rates of the low speed rotor (12,13,18) in the jet engines, means for providing second representative signals representative of rotation rates of the high speed rotor (14,15,17) in the jet engines; means (72,74-76,79,80) for providing threshold signals; and engine failure indicator means (120), characterised in that the apparatus further comprises equivalent signal means (92-95) coupled to receive the second representative signals for deriving equivalent signals representative of low speed rotor rotation rates of normally operating jet engines having high speed rotor rotation rates represented by the second representative signals; in that the difference means include first difference signal means (131) coupled to receive the first representative signals and the equivalent signals for providing first difference signals representative of differences between the first representative signals and the equivalent signals; and first detector means coupled to the threshold means and the first difference signal means for comparing the threshold signals and said first difference signals to provide first detected signals representative of values of the first difference signals relative to the threshold signals; the engine failure indicator means (120) being coupled to receive the detected signals for determinations of engine failures. Apparatus according to claim 1, characterised in that equivalent signal means includes scale factor means coupled to the representative signal means for scaling the second representative signals in accordance with a first preselected scale factor; constant value means for providing first constant value signals; and in that the difference means further include second difference signal means coupled to the scale factor means and to the constant value signal means for providing second difference signals representative of differences between the scaled second representative signals and the constant value signals to establish the equivalent signals; and limiter means coupled between the first and second difference signal means for providing the equivalent signals at values bound by preselected limits. Apparatus according to claim 1 characterised in that the threshold signal means comprises selection means coupled to receive the first representative signals for selecting one of the first representative signals having a value greater than values of all other first representative signals, thereby providing a first representative signal of maximum value; scale factor means coupled to the selection means for scaling the first representative signal of maximum value in accordance with a preselected scale factor, thereby providing a scaled maximum value signal; constant value means for providing a constant value signal; second difference signal means coupled to the scale factor means and the constant value means for providing second difference signals representative of differences between the scaled maximum value signal and the constant value signal; and limiter means coupled between the second difference signal means and the detector means for limiting the second difference signal between predetermined bounded values to establish a first threshold signal for the detector means. Apparatus according to claim 3, characterised in that the second representative signals include signals representative of rotation rates of the high speed rotor (14,15,17) in first and second jet engines; in that the equivalent signal means is coupled to receive the first and second jet engine high speed rotor rotation rate representative signals for providing equivalent low speed rotor rotation rate representative signals of normally operating first and second jet engines; in that the first difference signal means is coupled to receive the first representative signals of the first jet engine and the equivalent low speed rotor rotation rate representative signals of the first and second jet engines for providing the first difference signals comprising first equivalent difference signals representative of differences between the first representative signals of the first jet engine and the equivalent low speed rotation rate representative signals of the first jet engine, and second equivalent difference signals representative of differences between the first representative signals of the first jet engine and the equivalent low speed rotation rate representative signals of the second jet engine; in that the first detector means is coupled to receive the first and second equivalent difference signals for providing first detected equivalent signals representative of values of the first equivalent difference signals relative to the threshold signal and second detected equivalent signals representative of values of the second equivalent difference signals relative to the threshold signal and in that the apparatus further includes: polarity detector means coupled to receive the second equivalent difference signal for determining signal polarity of the second equivalent difference signals and providing polarity representative signals indicative of the determined signal polarity; and means coupled to the polarity detector means and to receive the first and second detected equivalent signals for providing engine operation indication signals to the engine failure indicator means (120). Apparatus according to claim 4, characterised in that the engine operation indication signal means includes a first AND gate (117) having first input means coupled to receive said polarity representative signals, second input means coupled to receive said second detected equivalent signals and output means coupled to the engine failure indicator means (120); and a second AND gate (118) having inverting input means (119) coupled to receive the polarity representative signals, non-inverting input means coupled to receive the second detected equivalent signals, and output means coupled to the engine failure indicator means (120). Apparatus according to any of the preceding claims, characterised in that it further includes means for providing signals representative of aircraft forward velocity; means coupled to receive the forward velocity representative signals for providing signals representative of windmilling rotation rates due to the forward velocities; second difference signal means (151) coupled to receive the first representative signals and the windmilling representative signals for providing second difference signals representative of differences between the first signals and the windmilling representative signals; and first polarity detector means (153) coupled between the second difference signal means and the engine failure indicator means (120) for receiving the second difference signals and for providing first polarity representative signals to the engine failure indicator means representative of polarities of the second difference signals. Apparatus in accordance with claim 6, characterised in that the windmilling rotation rate means includes first constant value signal means (104) for providing constant value signals; first scale factor means (102) coupled to the aircraft forward velocity signal means for providing the signals representative of aircraft forward velocity scaled in accordance with a preselected scale factor; sum means (103) coupled to the first scale factor means and to receive the constant value signals for providing signals representative of sums of the constant value signal and the scaled aircraft forward velocity representative signals; and first limiter means (105) coupled between the sum means and the second difference signal means for bounding the sum signals between preselected limits and providing bounded signals to the second difference signal means. Apparatus according to claim 7, characterised in that the first representative signals include signals representative of rotation rates of the low speed rotor in first and second jet engines; the first difference signal means is coupled to receive the low speed rotor representative signals of the first and second jet engines for providing third difference signals representative of differences between the low speed rotor representative signals of the first jet engine and the low speed rotor representative signals of the second jet engine; the threshold signals means provides a first threshold signal; the first detector means is coupled to receive the third difference signals and the first threshold signal for providing low speed rotor detected difference signals representative of relative values of the second difference signals and the first threshold signal; and in that the apparatus further includes; second polarity detector means coupled to receive the third difference signals for determining the signal polarity of the third difference signals and providing second polarity representative signals indicative of the determined signal polarity of the third difference signal; and engine operation indication signal means coupled to receive the second polarity representative signals and to receive the low speed rotor detected difference signals for providing low speed rotor data to the engine failure indicator means (120). Apparatus according to claim 8 characterised in that the second representative signals include signals representative of rotation rates of the high speed rotor (14,15,17) in first and second jet engines; the first difference signal means is coupled to receive the high speed rotor representative signals of the first and second jet engines for providing fourth difference signals representative of differences between the high speed rotor representative signals of the first jet engine and the high speed rotor representative signals of the second jet engine; the threshold signal means provides a second threshold signal; the first detector means is coupled to receive the fourth difference signals and the second threshold signal for providing high speed rotor detected difference signals representative of relative values of the third difference signals and the second threshold signal; the second polarity detector means is coupled to receive the fourth difference signals for determining signal polarity of the fourth difference signals and providing third polarity representative signals indicative of the signal polarity of the fourth difference signal; and engine operation indication signal means is coupled to receive the third polarity representative signals and to receive the high speed rotor detected difference signals for providing high speed rotor data to the engine failure indicator means (120). Apparatus according to claim 9, characterised in that the equivalent signal means includes second scale factor means coupled to the representative signal means for scaling the second representative signals in accordance with a first preselected scale factor to provide scaled second representative signals; second constant value means for providing second constant value signals; and third difference signal means coupled to the second scale factor means and the second constant value signal means for providing signals representative of differences between the scaled second representative signals and the second constant value signals to establish the equivalent signals; and second limiter means coupled between the first and third difference signal means for providing the equivalent signals at values bound by preselected limits. Apparatus according to claim 10, characterised in that the threshold signals means comprises selection means coupled to receive the first and second representative signals for selecting one of the first representative signals having a value greater than values of all other first representative signals and selecting one of the second representative signals having a value greater than values of all other second representative signals, thereby providing first and second representative signals of maximum value; third scale factor means coupled to the selection means for scaling the first representative signal of maximum value in accordance with a preselected scale factor, thereby providing a first scaled maximum value signal; second constant value means for providing a second constant value signal; fourth difference signal means coupled to the third scale factor means and the second constant value means for providing fifth difference signals representative of differences between the scaled first representative maximum value signal and the second constant value signal; third limiter means coupled between the fourth difference signal means and said first detector means for limiting the second difference signal between predetermined bounded values to establish a first threshold signal for the first detector means; fourth scale factor means coupled to the selection means for scaling the second representative signal of maximum value in accordance with a predetermined scale factor, thereby providing a second scaled maximum value signal; third constant value means for providing a third constant value signal; fifth difference signal means coupled to the third scale factor means and the third constant value means for providing sixth difference signals representative of differences between the scaled second representative maximum value signal and the third constant value signal; fourth limiter means coupled between said fifth difference signal means and the first detector means for limiting the sixth difference signals between preselected bounded values to establish a second threshold signal for the first detector means. Apparatus according to claim 11, characterised in that the equivalent signal means is coupled to receive the first and second jet engine high speed rotor rotation rate representative signals for providing equivalent low speed rotor rotation rate representative signals of normally operating first and second jet engines; the first difference signal means is coupled to receive the first representative signals of the first jet engine and the equivalent low speed rotor rotation rate representative signals of the first and second jet engines for providing the first difference signals comprising first equivalent difference signals representative of differences between the first representative signals of the first jet engine and the equivalent low speed rotation rate representative signals of the first jet engine, and second equivalent difference signals representative of differences between the first representative signals of the first jet engine and the equivalent low speed rotation rate representative signals of the second jet engine; the first detector means is coupled to receive the first and second equivalent difference signals for providing first detected equivalent signals representative of values of the first equivalent difference signals relative to the first threshold signal and second detected equivalent signals representative of values of the second equivalent difference signals relative to the first threshold signal; and in that the apparatus further includes third polarity detector means coupled to receive the second equivalent difference signals for determining the signal polarity of the second equivalent difference signals and providing third polarity representative signals indicative of the determined signal polarity to the engine operation indication signal means wherefrom engine operation indication signals are coupled to the failure indicator means (120). A jet engine testing apparatus in accordance with claim 12, characterised in that the engine operation indication means includes a first AND gate having first input means coupled to receive the third polarity representative signals, second input means coupled to receive the second equivalent difference signals and output means coupled to the engine failure indicator means (120); and a second AND gate having inverting input means coupled to receive the third polarity representative signals non-inverting input means coupled to receive the second equivalent difference signals, and output means coupled to the engine failure indicator means (120). Apparatus according to claim 1, characterised in that it further includes means (51) for providing third representative signals representative of forward velocities of the aircraft; sensor level detector means (33,36,52) having pre-established first, second and third threshold signal levels coupled to receive the first, second, and third representative signals for providing first, second, and third output signals when the first, second, and third representative signals respectively exceed the first, second, and third threshold signal levels; and means (38,42,43,53,54) coupled to the sensor level detector means for processing the first, second, and third output signals to determine operational status of the representative signal means. Apparatus according to claim 14, characterised in that the processing means include an exclusive OR gate (38) coupled to receive the first and second representative signals; a first AND gate (42) having a first input terminal coupled to the exclusive OR gate and a second input terminal coupled to receive the first representative signals; a second AND gate (43) having a first input terminal coupled to the exclusive OR gate and a second input terminal coupled to receive the second representative signals; and first and second latches (53,54) having output terminals coupled to the engine failure indicator means (120) and input terminals coupled respectively to the first and second AND gate, each further coupled to receive the third representative signals as latch enabling signals.

Inventors: ZWEIFEL, TERRY L.
Applicants: HONEYWELL INC.
Application year: 1986
Main classification: G06F15
IPC codes: G05D1, G01M15, G07C3, F02C9, F02K99