Internal combustion engine detonation indication system

Abstract: The invention relates to a detonation indication system for an internal combustion piston engine of an aircraft, which senses detonation by means of a sensor comprising a piezoelectric force transducer (12) installed under a the spark plug of each cylinder of the engine. The transducer produces an on-going combustion pressure-induced charge signal which is converted into a voltage signal by a charge converter (14). A comparator gating circuit (16) removes, from the voltage signal, signal components, associated with intake and exhaust strokes, which occur below a predetermined combustion pressure. A high-pass frequency filter (18) then removes signal components associated with low frequency changes in combustion pressure. The output from the filter (18) is applied to an RMS-to-direct current converter (20) which converts the output into a direct current voltage proportional to the RMS value of the filter output. The output from the converter (20) is applied to peak and hold circuitry (22) which holds the peak value of the converter output corresponding to the most severe detonation event which occurs during an adjustable reset time period. A digital display (24) is provided which receives signals from the circuit (22) corresponding to the peak values occurring during successive reset periods and provides a numerical read-out indicative of the severity of the associated detonation events.

Description: This invention relates generally to a detonation indication system, and more specifically to an onboard aircraft engine detonation indication system, which senses detonation through the use of combustion pressure sensing force transducers located under the spark plugs of each cylinder. The threat of imposing unleaded aviation gasoline regulations is causing great concern in the aviation industry, since most of the larger displacement piston engines of the existing fleet require 100 octane fuel. There is no economically acceptable alternate method of producing 100 octane fuel in the absence of tetra ethyl lead. The proposed changes in fuels will force a reduction of the octane levels which will greatly affect the performance of these large piston aircraft engines in their present configurations. The power to weight performance of the aircraft is, of course, crucial and the engines must operate at their peak performance levels with these anticipated octane reduction changes. The key means to restore power losses is the development of appropriate engine modifications and/or tuning changes. However, a tradeoff in developing engine modifications and tuning changes is the risk of detonation (knock) which can be destructive to an engine. It is anticipated that the lesser octane unleaded fuels of the future may require with potential revisions as to lifting recertification of a significant number of aircraft with potential revisions as to lifting capacity, load and engine performance, depending on the combined impact of fuel knock characteristics and engine modifications or tuning changes adopted. The present system was designed to perform ground and flight evaluations of new candidate unleaded aviation gasolines, and to facilitate the extensive recertification effort that lies ahead. Current methods of monitoring knock lack the precision required to evaluate new fuels, and place difficult monitoring and management demands on test pilots in the air. A detonation indication system with simplified management characteristics is needed, which can be quickly read without adjustments to determine engine knock for each cylinder of each engine. The instrument needs to detect not only intensity of the knock events but provide an indication of events frequency as well. One of the first aircraft engine analyzers was the Sperry system which is set forth in U.S. Patent No. 2,518,427. The Sperry system which incorporates an oscilloscope display, analyzed various other malfunctions, besides detonation such as defective valve train components, spark plugs, magnetos, incorrect ignition timing, mixture and various other engine irregularities. The Sperry system has been accepted as the standard by the regulatory agencies in the certification of aircraft, However, it is large, cumbersome and requires a substantial amount of accessory equipment which renders it difficult to use in flight for general aviation aircraft, and non-usable on the smaller members of this fleet. The Sperry system also requires substantial adjustment before it can be used, along with a technician to perform functional mode selections, and to view and adjust the oscilloscope which certainly could not be safely done by the test pilot alone. The AVCO Detonation Analyzer System was developed in more recent times. However, it also sensed vibration as in the Sperry system and presented the data on the screen of an oscilloscope. The AVCO system required a vibration pickup mounted on each cylinder and an engine mounted crankshaft encoder for referencing TDC (top dead center of each cylinder). The AVCO system demands a skilled operator to perform adjustments and considerable time of engine operation under knock conditions in the air, since the interpretation criteria is based on a count of knock events (light flashes) per minute. Other modern engine sensors in the non-aviation field have been developed for various automotive applications. U.S. Patent No. 4,450,811 belonging to the NGK Spark Plug Co., Ltd. is an example. The NGK controller is not a system to analyze the performance of an engine but rather a system for operating a fuel injected automotive engine which controlled the ignition timing and mixture to operate the fuel system of the vehicle. This system senses pressure in the cylinders through the use of piezoelectric washers on the head bolts. The NGK system requires measuring the crank angle (crankshaft encoder), which the present invention does not, and its purpose is somewhat different in that when it senses knock, it automatically retards the timing. Similar piezoelectric washers, as disclosed in the NGK patent above, have been used under spark plugs to sense knock in some Nissan automobiles for the same purpose as in the NGK system. There are automotive engine analyzers on the market today such as the AVL Indiskop and Norland systems which are stationary units used in laboratories for engine development and diagnostic purposes. These systems provide a visual indication of knock conditions by means of cylinder pressure traces displayed on a screen, along with numerous other engine parameters not embodied in this invention. U.S. Patent No. 5,062,294 discloses a ring-shaped sensor for detecting pressure and temperature in a cylinder of an internal combustion engine through the use of a piezoelectric transducer combined with a thermocouple positioned under the spark plug of the cylinder. This apparatus is intended to give a precise pressure wave measurement, and is not concerned with detonation. It is an object of the present invention to provide a detonation indication system, for example an onboard detonation indication system for aircraft piston engines, which is capable of detecting and measuring detonation in each cylinder with a numerical digital display. Another object of the present invention is to provide a detonation indication system which does not require the sensing of crank angle timing or encoder functions. A further object of the present invention is to provide a highly sensitive detonation detection and indication system for each cylinder which is capable of measuring both the minimum frequency of detonation events and their intensities. The invention provides a detonation indication system as defined in claim 1. The system embodying the present invention precisely detects detonation, and provides a measure of detonation intensity and minimum knock event frequencies in an aircraft internal combustion piston engine in flight. The detection system senses normal and abnormal combustion characteristics through piezoelectric force transducers installed under one spark plug in each cylinder, and renders a digital display which identifies isolated and repetitive detonation events in each cylinder. The pressure-induced signal is converted from an inline charge to a voltage. The voltage signal is then passed through a comparator gating means which segregates out the low pressure signals associated with intake and exhaust strokes that normally include severe vibratory components from the operation of the valve system. The comparator gating eliminates the need for shaft encoders which are required in the above-mentioned prior art systems. The signal then passes through a high-pass frequency filter which removes the normal compression and expansion cycle pressure signals and enhances the descrimination between normal and detonation combustion events. The filtered signal is then processed through an RMS (root mean square) to a DC converter which continuously computes the instantaneous square of the input signal from the high-pass filter, averages it, and takes the square root of the result to provide a DC voltage proportional to the RMS of the input. The DC signal is then processed through a peak and hold circuit which holds the peak signal for an adjustable reset-time period ranging between two and six seconds. Lastly, a digital display is provided for each cylinder which receives and displays a numerical value of, or corresponding to, the peak RMS signal. Reference will now be made to the accompanying drawings, in which:- Figure 1 is a block diagram of the various components of a detonation indication system embodying the present invention; Figure 2 includes time traces illustrating wave forms of various raw, filtered and gated signals in the circuits of the system; Figure 3 also depicts time traces illustrating the signal wave form changes of various components of the system; Figure 4 is a front elevational view of the control panel of the onboard aircraft detonation indication system; and Figure 5 is a partially sectional view, to an enlarged scale, of a modified form of sensor washer positioned under an aircraft spark plug. An onboard aircraft engine detonation indication system embodying the invention is generally identified by reference numeral 10. In Fig. 1, the system is shown in block diagram form for a single cylinder. Sensor 12, not shown in detail, is a quartz piezoelectric force transducer washer similar to those known in the automotive market and an example of which is shown in the above-mentioned NGK patent. The pressure sensors 12, shaped like washers with a wire connector and lead, are mounted under the spark plugs of each cylinder and torqued in place. An example of this is shown in Figure 5 with modified sensor 12', which will be described hereafter. The forces transmitted through the spark plugs to the sensors 12 during the combustion stroke in each cylinder generate an electrical charge signal with a time trace similar to that shown in Fig. 2(a) and Fig. 3(a). Located in the engine compartment in close proximity to each sensor 12 is an inline charge converter 14 which converts the signal charge to voltage as illustrated by the time trace in Fig. 3(b). In Fig. 2(a) a raw time trace of two consecutive combustion cycles which occur within the engine cylinder is illustrated, showing a normal combustion event on the right and a detonation combustion event on the left. With detonation, a series of high-frequency spikes develop commencing near top dead center, as graphically shown on the left, which are superimposed on the normal low-frequency compression and expansion cycle pressures. Fig. 2(a) also illustrates a variety of vibratory components associated with the opening and closing of the intake and exhaust valves which take place at low combustion chamber pressures, all of which are of no concern to the detection and intensity assessment of detonation. The raw voltage signals shown in Fig. 2(a) are gated at a threshold level as indicated in Fig. 2(b) wherein all signals below the threshold level are excluded by a threshold circuit such as comparator gating means 16. The gated signal is then passed through a high-pass frequency filter 18 to remove the low-frequency cycle pressures, which leaves a signal with a time trace as illustrated in Fig. 2(d). Thus, the high-pass filter 18 takes advantage of the emergence of the high-frequency spikes associated with abnormal combustion or detonation events, and, as will be apparent from a comparison of the left and right-hand components of the signal in Fig. 2(d), enhances the discrimination between detonation and normal combustion events. The high-pass filter 18 may, for example, be an eight pole filter having a combined roll-off rate of approximately 48dB/Octave. In the absence of the comparator gating means 16, the high-pass filter 18 would emit a signal as shown in Fig. 2(c). This signal, with all of the extraneous other relatively high-frequency signal spikes, would obviously overshadow the pressure signals derived strictly from normal or detonation combustion processes. In the prior art, this problem was dealt with through the use of a shaft encoder which looked at the signal only during a portion of the compression stroke and combustion stroke of a combustion cycle. With the comparator gating means 16, this is not necessary, and the overall indication system becomes simpler to install and more reliable in operation. The RMS values of the signal emitted from the high-pass frequency filter as illustrated in Figs. 2(d) and 3(d) are converted to a DC voltage by circuit 20 as illustrated by the time trace in Fig. 3(e). The peak and hold circuit 22 takes the signal from the circuit 20, while searching for, and holding, the value of the peak or maximum signal which occurs during each of a succession of reset time periods. The output from the peak and hold circuit for a plurality of successive reset periods is indicated by the time trace in Fig. 3(f). At the end of each reset period, the DC level corresponding to the peak signal for that period is output and maintained until the next reset period is completed, at which time it is replaced by the new (higher or lower) DC output level. This single peak signal represents the most severe normal or detonation combustion event which occurs during that reset time period. The reset time period is adjustable, and may be adjusted between approximately two and six seconds, for example to 5.5 seconds which would capture approximately 100 combustion cycles at 2200 engine RPM and approximately 124 combustion cycles at 2700 RPM, which is the range of engine speeds covering the most significant spectrum of detonation investigations on current generation aircraft engines. The digital display 24 translates the peak and hold output signal to an effectively real-time digital reading, as seen in Figs. 4 and 3. During calibration, the output level of the high-pass filter 18, and/or a reference point of the digital display 24, may be adjusted so as to indicate, or detect, the transition from normal combustion to knocking, and to illuminate an indicator light 26 as shown in Figs. 3 and 4 which flashes for each successive reset time period that experiences at least one detonation event. For example, the filter output level/display reference detonation point could be set to digitally display a numerical value of 100 at the onset of detonation, and this level will also cause the light 26 to flash. During operation of the system, the numerical value displayed will be indicative of, or proportional to, the output of the peak and hold circuit, i.e. the RMS signal corresponding to the most severe detonation event experienced during each reset time period, or the most severe normal combustion event in the absence of detonation. Under normal combustion conditions, this numerical value or combustion severity number may, for example, float up to levels of 35 to 50, rising to 100 at the onset of detonation, and increasing above 100 as the severity of detonation increases during each reset time, for example to or above a level of 600. Thus, the peak and hold circuit, working in conjunction with the RMS computation and conversion to direct current functions, provides DC voltage outputs that represent the most severe combustion signals experienced over a finite number of combustion cycles dependent upon engine speed. Fig. 4 illustrates the control panel 30 of the detonation indication unit without the pressure sensors 12 or the charge converters 14, both of which are located on the engine, while the indication unit 30 is positioned in the cockpit adjacent the pilot or observer. Due to the simplified digital read out of each cylinder and its flashing light, the pilot of the aircraft can utilize the detection system 10 while changing the operating conditions of the engine in flight. Various factors bring on detonation or knock including reduction in fuel octane ratings, ignition timing changes, induction air manifold pressure, induction air temperatures and RPM. While the control panel 30 will indicate each reset period with at least one detonation event for each cylinder, it is sometimes desirable to have the option of utilizing an oscilloscope to view the raw and filtered signal traces for a particular cylinder. To achieve this, an oscilloscope 34 shown symbolically in Fig. 3(c) can be connected to the control panel terminals 36 and 38 in Figure 4. The system for each cylinder includes a gain adjustment 44, which after being initially set generally requires no further adjustment. Fig. 4 only illustrates three electrical leads 40 connected to the control panel 30, however, when in use, the other leads would be in place depending upon the number of cylinders per engine. The RMS ports 42, in Fig. 4, may be connected to a data acquisition means 46 symbolically shown in Fig. 3(f), to record combustion-intensity digital indications 24, concurrently with other engine and flight test data, and to automatically activate the data acquisition means at the onset of detonation. Figure 5 illustrates a modified form of sensor washer 12' symbolically shown positioned under spark plug 50 in an aircraft cylinder head 52. On top of the sensor 12' is a conventional steel washer 58 which exerts a compressive load on sensor 12' when the plug 50 is torqued in place. Located within sensor 12' are two different sensors, one being a piezoelectric force transducer 68 as previously described and, secondly, a conventional thermocouple 66 for sensing cylinder head temperature, both of which are contained within a common sheath. Exiting sensor 12' on the right side is the piezoelectric lead 54, and on the left side, the thermocouple lead 56. Lead 54 connects to input terminal 60 shown in Figure 4 while lead 56 connects to CHT indicator 62 shown in Figure 3(a). In typical aircraft detonation-testing, cylinder head temperatures are recorded by separate sensors on each cylinder head. With the optional data acquisition means 46 in use, there is no calculation and recording of data until the initial detonation event, at which time it begins to record data automatically including number of events, combustion severity and additional data such as CHT if desired. Means 46 can also be activated manually. The system 10 is basically free of adjustments and merely requires positioning a washer-like pressure sensing transducer or sensor 12 under each of the cylinder's spark plugs and to connect the leads 40 to the control panel 30. The pressure sensors 12 not only indicate the onset of detonation but they also measure the intensity of the detonation in the digital display 24. The system 10 is very accurate and can measure the onset of detonation within 2.54 mm (0.1 inches) of induction air manifold pressure. One of the intended uses of the system 10 is with a fuel system which blends two different octane fuels during flight to adjust the octane rating of the blended fuel stream supplying the engine.

Claims: A detonation indication system for an internal combustion piston engine which senses pressure within a cylinder of the engine, including: sensor means (12) comprising piezoelectric force transducer means, associated with a spark plug (50) of the cylinder to produce an on-going pressure-induced charge signal representative of the combustion cycles occurring within the cylinder; and circuitry for processing the charge signal to derive an output indicative of detonation occurring within the cylinder; characterised by: inline charge-to-voltage conversion means (14) for converting the charge signal from the transducer means into a voltage signal; comparator gating means (16) to remove, from the voltage signal, low pressure signal components associated with intake and exhaust strokes; high-pass frequency filter means (18) to remove, from the voltage signal, signal components associated with low frequency changes in combustion pressure; RMS-to-direct current converter means (20) to convert the gated and filtered voltage signal into a direct current output voltage signal proportional to the RMS value of the voltage signal; peak and hold circuitry means (22) which selects and holds, for each of a series of adjustable reset time periods, the peak value of the output signal from the RMS-to-direct current converter means which is indicative of the severity of a combustion event; and display means (24) to receive signals corresponding to the peak values occurring during successive reset periods and provide a display indicative of detonation events. A detonation indication system as claimed in claim 1, wherein the display means comprises digital display means (24) operable to display a numerical value indicative of the transition from normal combustion to detonation. A detonation indication system as claimed in claim 1 or 2, wherein the display means comprises digital display means (24) operable to display a numerical value indicative of the maximum severity of detonation. A detonation indication system as claimed in claim 1, 2 or 3, wherein the force transducer means is of a quartz type. A detonation indication system as claimed in any preceding claim, wherein the sensor means (12) comprises a ring-shaped sensor installed under the spark plug (50), and includes a piezoelectric force transducer for sensing pressure within the cylinder, and a thermocouple for sensing cylinder head temperature, both contained within a common shroud.

Inventors: CHEEK JAMES, GONZALEZ CESAR, CHEEK, JAMES, GONZALEZ, CESAR
Applicants: CESSNA AIRCRAFT CO, CESSNA AIRCRAFT COMPANY
Application year: 1995
Main classification: G01L23
IPC codes: G01L23