Sensing system

Abstract

To accurately estimate the attitude or motion state of a flying object.SOLUTION: A sensing system comprises: a first reference pressure room within a flying object which is a space connected via a first communication passage to a plurality of holes for measuring the reference pressure that are provided on the surface of a tip on the front side of the flying object; a second reference pressure room within the flying object which is a space connected via a second communication passage to the plurality of holes for measuring pressure that are provided on the surface of the tip; a plurality of differential pressure sensors that are provided in the second communication passage and detect the differential pressure between the pressure on the side of the holes for measuring pressure and the pressure on the side of the second reference pressure room; an absolute pressure sensor that is provided in the second reference pressure room and detects an absolute pressure; and an estimation unit for estimating the attitude or motion state of the flying object on the basis of the output of the differential pressure sensor and the output of the absolute pressure sensor.SELECTED DRAWING: Figure 2

Description

[Technical Field]

[0001] This invention relates to a sensing system. [Background technology]

[0002] In aircraft, Air Data Sensing (ADS) systems using pitot tubes and vanes are used to estimate air data (such as airspeed and aerodynamic attitude). However, in the case of hypersonic flight, such as in spaceplanes, the sharp shape of ADS systems makes them difficult to use because the tip is exposed to high temperatures due to the effects of shock waves. Therefore, an air data estimation method using a Flush Air Data Sensing (FADS) system, which has a blunt-shaped nose surface with numerous pressure holes, is known (for example, Non-Patent Document 1). [Prior art documents] [Non-patent literature]

[0003] [Non-Patent Document 1] Whitmore, SA, Cobleigh, BR and Hearing, EA, Design and Calibration of the X-33 Flush Airdata Sensing (FADS) System, NASA / TM-1998-206540, 1998. [Overview of the Initiative] [Problems that the invention aims to solve]

[0004] However, conventional FADS systems have a problem in that estimation accuracy decreases in the low dynamic pressure region immediately after liftoff and near the apex of orbit. This is because pressure sensors with a large range that can measure the pressure of each hole during high dynamic pressure flight cannot achieve the measurement accuracy necessary to capture the differential pressure between holes during low dynamic pressure. It is necessary to switch between pressure sensors with multiple pressure ranges, but this increases costs.

[0005] Therefore, it is conceivable to provide a reference pressure chamber that guides the average pressure on the nose surface and narrow the range required for the pressure sensor by using differential pressure measurement between each hole and the reference pressure chamber, thereby improving the estimation accuracy.

[0006] At this time, since the pressure on the nose surface is calculated using the difference from the reference pressure measured by each differential pressure gauge, a design of the reference pressure chamber is required such that there is no pressure difference between multiple reference pressure measurement points.

[0007] The present invention has been made in view of the above circumstances, and an object thereof is to provide a sensing system capable of accurately estimating the attitude state or motion state of a flying object.

Means for Solving the Problems

[0008] The sensing system according to the present invention includes a plurality of reference pressure measurement holes provided on the surface of the tip portion on the front side of the flying object, a first reference pressure chamber in the flying object that is a space connected via a first communication path, a plurality of pressure measurement holes provided on the surface of the tip portion, a second reference pressure chamber in the flying object that is a space connected via a second communication path, a plurality of differential pressure sensors provided in the second communication path for detecting the differential pressure between the pressure on the pressure measurement hole side and the pressure on the second reference pressure chamber side, an absolute pressure sensor provided in the second reference pressure chamber for detecting the absolute pressure, and an estimation unit for estimating the attitude state or motion state of the flying object based on the output of the differential pressure sensors and the output of the absolute pressure sensor.

[0009] According to the present invention, the first reference pressure chamber in the flying object is connected to a plurality of reference pressure measurement holes provided on the surface of the tip portion on the front side of the flying object, and the second reference pressure chamber in the flying object is connected to a plurality of pressure measurement holes provided on the surface of the tip portion. Further, the differential pressure sensors are provided in the second communication path connecting the pressure measurement holes and the second reference pressure chamber, and the absolute pressure sensor is provided in the second reference pressure chamber.

[0010] Then, the estimation unit estimates the attitude state or the motion state of the flying object based on the output of the differential pressure sensor and the output of the absolute pressure sensor.

[0011] In this way, by providing the first reference pressure chamber and the second reference pressure chamber and using the output of the absolute pressure sensor provided in the second reference pressure chamber and the output of the differential pressure sensor provided in the second communication path, the attitude state or the motion state of the flying object can be accurately estimated.

[0012] The third communication path connecting the first reference pressure chamber and the second reference pressure chamber according to the present invention can be formed at the center of the surface facing the surface on which the first communication path is formed. Thereby, the influence on the second reference pressure chamber due to the gas flow in the first reference pressure chamber can be suppressed.

[0013] The estimation unit according to the present invention calculates the pressure of the pressure measurement hole based on the output of the differential pressure sensor and the output of the absolute pressure sensor, and estimates the attitude state or the motion state of the flying object based on the pressure of the pressure measurement hole. Thereby, the pressure of the pressure measurement hole can be accurately calculated, and the attitude state or the motion state of the flying object can be accurately estimated.

[0014] The sensing system according to the above invention includes a plurality of the absolute pressure sensors, the plurality of absolute pressure sensors are provided in a plurality of absolute pressure measurement holes formed in the second reference pressure chamber, and further includes an abnormality determination unit that determines an abnormal state of each of the plurality of differential pressure sensors and the plurality of absolute pressure sensors. The estimation unit calculates the pressure of the second reference pressure chamber based on the output of the absolute pressure sensor that is not determined to be in an abnormal state, and estimates the attitude state or the motion state of the flying object based on the output of the differential pressure sensor that is not determined to be in an abnormal state and the pressure of the second reference pressure chamber. Thereby, even if an abnormality occurs in the pressure measurement hole or the absolute pressure measurement hole, the attitude state or the motion state of the flying object can be accurately estimated.

Effects of the Invention

[0015] As described above, according to the sensing system of the present invention, by providing a first reference pressure chamber and a second reference pressure chamber, and using the output of an absolute pressure sensor provided in the second reference pressure chamber and the output of a differential pressure sensor provided in the second communication passage, the attitude state or motion state of an aircraft can be estimated with high accuracy. [Brief explanation of the drawing]

[0016] [Figure 1] (A) A front view of the front end of an aircraft according to an embodiment of the present invention, and (B) a side view of the front end. [Figure 2] This is a schematic diagram showing the configuration of a sensing system according to an embodiment of the present invention. [Figure 3A] This is a cross-sectional view of line AA in Figure 1(A). [Figure 3B] This is a cross-sectional view of line BB in Figure 1(A). [Figure 4] This is a block diagram showing an estimation device for a sensing system according to an embodiment of the present invention. [Figure 5] This flowchart shows the contents of the estimation process performed by the estimation device of the sensing system according to an embodiment of the present invention. [Figure 6] This graph shows the flight profile of the winged rocket experimental aircraft in the embodiment. [Figure 7] This graph shows the estimation error obtained by absolute pressure measurement. [Figure 8] This graph shows the estimation error obtained by differential pressure measurement. [Modes for carrying out the invention]

[0017] Embodiments of the present invention will be described in detail below with reference to the drawings.

[0018] <Summary of Embodiments of the Invention> Since the pressure on the front surface of the aircraft is calculated using the difference between the pressure measured by each differential pressure gauge and the reference pressure, the design of the reference pressure chamber is required so that no pressure differences occur between multiple reference pressure measurement points.

[0019] In the embodiment of the present invention, the shape of the reference pressure chamber inside the tip is designed to suppress the flow of gas inside in order to satisfy this requirement, thereby accurately estimating the pressure in the pressure measurement hole and improving the accuracy of air data estimation.

[0020] <Configuration of the sensing system according to the embodiment of the present invention> As shown in Figure 1, the sensing system according to an embodiment of the present invention is a differential pressure flash-type air data sensing system. In the sensing system according to an embodiment of the present invention, a plurality of reference pressure measuring holes r1 to r8 and a plurality of pressure measuring holes 1 to 9 are formed on the surface of the front tip portion 100F of the aircraft.

[0021] Figure 1(A) is a front view of the front tip 100F of the aircraft, showing an example in which eight reference pressure measuring holes r1 to r8 and nine pressure measuring holes 1 to 9 are arranged radially in the aircraft's housing. It also shows an example in a front view in which four reference pressure measuring holes r1 to r4 are arranged at 90° intervals on the inner circle and four reference pressure measuring holes r5 to r8 are arranged at 90° intervals on the outer circle. Furthermore, it shows an example in a front view in which pressure measuring hole 1 is located at the center of the circle, four pressure measuring holes 2 to 5 are arranged at 90° intervals on the inner circle and four pressure measuring holes 6 to 9 are arranged at 90° intervals on the outer circle.

[0022] Furthermore, as shown in Figure 1(B), an example is shown in which five pressure measuring holes 1, 2, 4, 6, and 8 are arranged at 15° intervals from the center of a circle approximating the curve of the surface of the tip portion 100F in a side view.

[0023] Figure 2 also shows a schematic diagram of the sensing system inside the front tip 100F of the aircraft. The sensing system comprises a first reference pressure chamber 10, a second reference pressure chamber 20, a plurality of differential pressure sensors 51-59, and a plurality of absolute pressure sensors 61-64.

[0024] The first reference pressure chamber 10 is a space within the aircraft connected to a plurality of reference pressure measuring holes r1 to r8 via first communication passages 31 to 38. The second reference pressure chamber 20 is a space within the aircraft connected to a plurality of pressure measuring holes 1 to 9 via second communication passages 41 to 49.

[0025] The differential pressure sensors 51-59 are located in the second communication passages 41-49 and detect the differential pressure between the pressure on the side of the pressure measuring holes 1-9 and the reference pressure on the opposite side.

[0026] Multiple absolute pressure sensors 61-64 are installed in the second reference pressure chamber 20 and detect absolute pressure.

[0027] Third connecting passages 81-83, which connect the first reference pressure chamber 10 and the second reference pressure chamber 20, are formed in the center of the surface opposite to the surface on which the first connecting passages 31-38 are formed.

[0028] Multiple absolute pressure sensors 61-64 are provided in multiple absolute pressure measuring holes 91-94 formed in the second reference pressure chamber 20.

[0029] Multiple reference pressure measuring holes r1 to r8 are connected to the first reference pressure chamber 10, thereby determining the average pressure on the surface of the tip portion 100F.

[0030] In this embodiment, by reducing the differential pressure between the second reference pressure chamber 20 and the pressure measurement holes 1-9 across a wide range of dynamic pressures, Mach numbers, and aerodynamic attitudes, the required range for the differential pressure sensors 51-59 is narrowed, thereby improving the accuracy of pressure estimation at pressure measurement holes 1-9. Note that using different reference pressures when calculating the pressure at each of the pressure measurement holes 1-9 will degrade the accuracy of pressure estimation; therefore, a design is required to ensure that there is no pressure difference between the reference pressure measurement points of the differential pressure sensors 51-59.

[0031] Therefore, the sensing system of this embodiment is equipped with a first reference pressure chamber 10 and a second reference pressure chamber 20, and the reference pressure chamber has a two-stage structure. When a pressure difference exists between the multiple reference pressure measuring holes r1 to r8, a gas flow occurs inside the first reference pressure chamber 10, but to prevent this from affecting pressure measurement, the second reference pressure chamber 20 is provided, which is connected to differential pressure sensors 51 to 59 and absolute pressure sensors 61 to 64. In order to reduce the pressure difference between the three third connecting passages 81 to 83 that connect the first reference pressure chamber 10 and the second reference pressure chamber 20 and to suppress the occurrence of pressure unevenness inside the second reference pressure chamber 20, the three third connecting passages 81 to 83 are formed in the center of the surface facing the surface on which the first connecting passages 31 to 38 are formed.

[0032] Figure 3A shows a cross-sectional view of the front tip 100F of the aircraft 100 along line AA in Figure 1(A). Figure 3B shows a cross-sectional view of the front tip 100F of the aircraft 100 along line BB in Figure 1(A). Figures 3A and 3B show an example in which a first reference pressure chamber 10 is provided inside the housing, and a second reference pressure chamber 20 is provided behind the first reference pressure chamber 10. Figures 3A and 3B show how the first reference pressure chamber 10 is connected to the reference pressure measuring holes r2, r4, r6, and r8 via the first connecting passages 32, 34, 36, and 38, and how the first reference pressure chamber 10 is not connected to the pressure measuring holes 1, 2, 4, 6, and 8.

[0033] Furthermore, the sensing system includes an estimation device 70 that estimates the attitude or motion state of the aircraft based on the outputs of differential pressure sensors 51-59 and absolute pressure sensors 61-64.

[0034] The estimation device 70 can be configured as a computer including a CPU (Central Processing Unit), RAM (Random Access Memory), and ROM (Read Only Memory) which stores programs and various data for executing the estimation processing routine described later. Functionally, the estimation device 70 includes an acquisition unit 72 and an estimation unit 76, as shown in Figure 4.

[0035] Here, a method for estimating the attitude and motion states of the flying object by the estimation device 70 will be described.

[0036] In this embodiment, absolute pressure measurement and differential pressure measurement are used in combination to estimate the pressures of the pressure measurement holes 1 to 9. In advance, the necessary calibration coefficients are obtained from experimental data such as wind tunnel tests.

[0037] During the flight of the flying object, when estimating the pressures of the pressure measurement holes 1 to 9, first, as inputs, the differential pressures Δp between the second reference pressure chamber 20 side and the pressure measurement holes 1 to 9 sides measured by the differential pressure sensors 51 to 59 i and the absolute pressures p measured by the absolute pressure sensors 61 to 64 j (r), the arrangements (front view angle φ i , side view angle λ i ) of each of the pressure measurement holes 1 to 9 and the reference pressure measurement holes r1 to r8 are acquired. Note that the arrangements (front view angle φ i , side view angle λ i ) of each of the pressure measurement holes 1 to 9 and the reference pressure measurement holes r1 to r8 are known as pre-designed values.

[0038] Next, the reference pressure p(r) of the second reference pressure chamber 20 is determined from the four absolute pressures p j (r). The pressures p of each of the pressure measurement holes 1 to 9 are obtained from the sum of the differential pressure Δp i and the reference pressure p(r). i

[0039] After the pressures p of the pressure measurement holes 1 to 9 i are calculated, as described below, the Triples method (Non-Patent Document 1) is applied to estimate the attitude and motion states of the flying object.

[0040] First, the local angle of attack α e is obtained from a plurality of combinations in which three holes are selected from the pressure measurement holes 1, 2, 4, 6, 8 located on the vertical line in the front view. Also, the local sideslip angle β i is calculated from the obtained p e and α e ​i , α e , β e From the previously obtained calibration coefficient ε, the Mach number M and dynamic pressure q are calculated. c , static pressure q ∞ We estimate the local angle of attack α. e and local sideslip angle β e The angle of attack correction value δα and the sideslip angle correction value δβ are calculated from this. Finally, local angle of attack α e and local sideslip angle β e From the angle of attack correction value δα and the sideslip angle correction value δβ, the angle of attack α and sideslip angle β relative to the main stream are obtained. e We estimate this.

[0041] Furthermore, Non-Patent Document 1 above applies the Triples method to a FADS having six pressure measurement holes. In this embodiment, by providing nine redundant pressure measurement holes 1 to 9, the pressure gauge achieves 1) fault tolerance (it can be used without deterioration of pressure estimation accuracy) and 2) fault detection (a fault flag f that warns of deterioration of estimation accuracy due to multiple faults). fail This achieves fault tolerance (output of ). Furthermore, although the accuracy of pressure estimation under low dynamic pressure has been improved by using a differential pressure sensor in combination, it cannot handle extremely low dynamic pressure, so it self-detects that the estimation accuracy cannot be guaranteed due to the decrease in dynamic pressure and sets an accuracy reduction flag f low It has a notification function.

[0042] As described above, in this embodiment, the acquisition unit 72 acquires the outputs of the differential pressure sensors 51 to 59 and the absolute pressure sensors 61 to 64.

[0043] The estimation unit 76 determines the abnormal state of each of the differential pressure sensors 51-59 and absolute pressure sensors 61-64. For example, by comparing the outputs of absolute pressure sensors 61-64, if the absolute value of the difference between it and the output of another absolute pressure sensor is greater than or equal to a threshold, it is determined that the absolute pressure is in an abnormal state. Then, angle of attack estimation and sideslip angle estimation are performed multiple times using the output of one of the differential pressure sensors, and if there is a difference of more than a threshold between the multiple estimated values, the differential pressure sensor that is commonly used for outlier estimation is determined to be in an abnormal state.

[0044] The estimation unit 76 generates a fault flag f according to the result of the abnormal condition determination. fail It outputs the following. In addition, the estimation unit 76 self-determines that the estimation accuracy cannot be guaranteed due to the decrease in dynamic pressure and outputs an accuracy reduction flag f low Outputs.

[0045] The estimation unit 76 calculates the pressure in pressure measuring holes 1 to 9 based on the outputs of differential pressure sensors 51 to 59 and absolute pressure sensors 61 to 64, and estimates the attitude state or motion state of the aircraft based on the pressure in pressure measuring holes 1 to 9.

[0046] Specifically, the estimation unit 76 calculates the pressure in the second reference pressure chamber 20 based on the output of the absolute pressure sensors among the multiple absolute pressure measuring holes 91 to 94 that have not been determined to be in an abnormal state.

[0047] Furthermore, the estimation unit 76 estimates the attitude state of the aircraft, specifically the angle of attack α and the sideslip angle β relative to the main stream, based on the output of the differential pressure sensors among the pressure measurement holes 1 to 9 that are not determined to be in an abnormal state, and the pressure in the second reference pressure chamber 20. It also estimates the motion state, specifically the Mach number M and the dynamic pressure q. c , static pressure q ∞ We estimate this.

[0048] <Operation of the sensing system according to the embodiment of the present invention> First, obtain the necessary calibration coefficient ε from experimental data such as wind tunnel tests.

[0049] Then, while the aircraft is in flight, differential pressure sensors 51-59, located in the second communication passages 41-49 connected to the multiple pressure measuring holes 1-9, output differential pressure, and absolute pressure sensors 61-64, located in the multiple absolute pressure measuring holes 91-94 formed in the second reference pressure chamber 20, output absolute pressure. At this time, the estimation process shown in Figure 5 is performed by the estimation device 70.

[0050] First, in step S100, the acquisition unit 72 acquires the outputs of the differential pressure sensors 51-59 and the absolute pressure sensors 61-64.

[0051] In step S101, the estimation unit 76 determines the abnormal state of each of the differential pressure sensors 51-59 and absolute pressure sensors 61-64.

[0052] In step S102, the estimation unit 76 calculates the pressure in the second reference pressure chamber 20 based on the output of the absolute pressure sensors among the multiple absolute pressure measuring holes 91 to 94 that have not been determined to be in an abnormal state.

[0053] In step S104, the estimation unit 76 determines the pressure p of the pressure measurement holes that are not determined to be in an abnormal state, based on the output of the differential pressure sensor among the absolute pressure measurement holes 1 to 9 that are not determined to be in an abnormal state, and the pressure in the second reference pressure chamber 20. i Calculate.

[0054] In step S106, the estimation unit 76 determines the local angle of attack α based on the pressure in the pressure measuring hole corresponding to the differential pressure sensor that has not been determined to be in an abnormal state. e We seek.

[0055] In step S108, the estimation unit 76 determines the pressure p of the pressure measuring port based on the pressure of the pressure measuring port corresponding to the differential pressure sensor that has not been determined to be in an abnormal state. i and local angle of attack α e From the local sideslip angle β e Calculate.

[0056] In step S110, the estimation unit 76 calculates the pressure p of the pressure measuring hole corresponding to the differential pressure sensor that has not been determined to be in an abnormal state. i , local angle of attack α e , local sideslip angle β e Then, using the previously obtained calibration coefficient ε, the Mach number M and dynamic pressure q are calculated. c , static pressure q ∞ We estimate this.

[0057] In step S112, the estimation unit 76 determines the local angle of attack α e and local sideslip angle β e From the angle of attack correction value δα and the sideslip angle correction value δβ, the angle of attack α and sideslip angle β relative to the main stream are obtained. e We estimate this.

[0058] In step S114, the estimation unit 76 determines a fault flag f according to the result of the abnormal condition determination. fail The estimation unit 76 also determines that the estimation accuracy cannot be guaranteed due to a decrease in dynamic pressure and sets an accuracy reduction flag f low To decide.

[0059] In step S116, the estimation device 70 determines the angle of attack α and the sideslip angle β relative to the main stream, the Mach number M, and the dynamic pressure q. c , static pressure q ∞ , failure flag f fail , accuracy reduction flag f low Outputs.

[0060] <Examples> To demonstrate that the sensing system described in the above embodiment has superior estimation accuracy compared to conventional FADS, the flight profile (Figure 6) of the winged rocket experimental aircraft WIRES#015 (Non-Patent Literature 2) was used as the analysis condition, and the estimation error was analyzed assuming a 3σ error (Table 1) for the absolute pressure sensor and differential pressure sensor.

[0061] [Non-patent document 2] Fukushima, D., Yonemoto, K., Fujikawa, T., Matsukami, T., Otsuki, T., Kitazono, Y., Koshida, Y., Murakami, M., Watanabe, T., and Morito, T., Design and Development Status of Experiment Winged Rocket WIRES#015, 33rd International Symposium on Space Technology and Science, 2022.

[0062] [Table 1]

[0063] Figures 7 and 8 show the results of error analysis performed on air data for angle of attack, sideslip angle, Mach number, flight dynamic pressure, and atmospheric static pressure. Figure 7 is a graph showing the estimation error using absolute pressure measurement, which is a conventional FADS. Figure 8 is a graph showing the estimation error using differential pressure measurement, as described in the above embodiment. As a result, it can be seen that estimation using differential pressure measurement, as described in the above embodiment, suppresses measurement errors of air data to an extremely small degree compared to conventional FADS.

[0064] As described above, according to the sensing system of the embodiment of the present invention, by providing a first reference pressure chamber and a second reference pressure chamber, and using the output of an absolute pressure sensor provided in the second reference pressure chamber and the output of a differential pressure sensor provided in the second communication passage, the attitude state or motion state of an aircraft can be estimated with high accuracy.

[0065] Furthermore, by forming a third connecting passage between the first and second reference pressure chambers in the center of the surface opposite to the surface where the first connecting passage is formed, the influence of the gas flow in the first reference pressure chamber on the second reference pressure chamber can be suppressed.

[0066] Furthermore, by calculating the pressure at the pressure measurement port based on the output of the differential pressure sensor and the absolute pressure sensor, the pressure at the pressure measurement port can be accurately calculated even at low dynamic pressures, and the attitude or motion state of the aircraft can be accurately estimated.

[0067] Furthermore, by determining the abnormal state of each of the multiple pressure measuring holes and the multiple absolute pressure measuring holes, calculating the pressure in the second reference pressure chamber based on the output of the absolute pressure sensor of the absolute pressure measuring hole that has not been determined to be in an abnormal state, and estimating the attitude or motion state of the aircraft based on the output of the differential pressure sensor of the pressure measuring hole that has not been determined to be in an abnormal state and the pressure in the second reference pressure chamber, the attitude or motion state of the aircraft can be estimated with high accuracy even if an abnormality occurs in the pressure measuring hole or absolute pressure measuring hole.

[0068] Furthermore, by using a two-stage structure for the reference pressure chamber in FADS, which employs differential pressure measurement to achieve high-precision air data estimation across a wide dynamic pressure range, the inflow of airflow into the second reference pressure chamber, where the reference pressure is measured, can be suppressed. In addition, error analysis along the flight profile of the experimental aircraft confirmed that FADS employing differential pressure measurement improves the accuracy of estimating the attitude or motion state of the aircraft.

[0069] In the above embodiment, the example described was one in which eight reference pressure measuring holes and nine pressure measuring holes are formed on the surface of the front tip of the aircraft, but the invention is not limited to this. The number of reference pressure measuring holes may be other than eight, and the number of pressure measuring holes may be other than nine. The arrangement of the reference pressure measuring holes and pressure measuring holes may be other than the arrangement shown in Figure 1 above. [Explanation of Symbols]

[0070] 1-9 Pressure measurement holes 10. First reference pressure chamber 20 Second reference pressure chamber 31~38 1st communication passage 41~49 2nd communication passage 51-59 Differential pressure sensor 61-64 Absolute pressure sensor 70 Estimation device 72 Acquisition Department 76 Estimation part 81~83 3rd passageway 91-94 Holes for absolute pressure measurement 100 flying objects 100F Tip r1~r8 Holes for measuring reference pressure

Claims

1. A first reference pressure chamber within the aircraft, which is a space connected via a first connecting passage to a plurality of reference pressure measuring holes provided on the surface of the front tip of the aircraft, The second reference pressure chamber within the aircraft body is a space connected via a second connecting passage to a plurality of pressure measuring holes provided on the surface of the tip portion, Multiple differential pressure sensors are provided in the second communication passage to detect the differential pressure between the pressure on the pressure measuring hole side and the pressure on the second reference pressure chamber side, An absolute pressure sensor for detecting absolute pressure is provided in the second reference pressure chamber, An estimation unit that estimates the attitude state or motion state of the aircraft based on the output of the differential pressure sensor and the output of the absolute pressure sensor, Includes, A sensing system in which a third connecting passage connecting the first reference pressure chamber and the second reference pressure chamber is formed in the center of a surface facing the surface on which the first connecting passage is formed.

2. The sensing system according to claim 1, wherein the estimation unit calculates the pressure in the pressure measuring hole based on the output of the differential pressure sensor and the output of the absolute pressure sensor, and estimates the attitude state or motion state of the aircraft based on the pressure in the pressure measuring hole.

3. The system includes multiple absolute pressure sensors, The plurality of absolute pressure sensors are provided in a plurality of absolute pressure measuring holes formed in the second reference pressure chamber. The estimation unit, The abnormal state of each of the plurality of differential pressure sensors and the plurality of absolute pressure sensors is determined, Based on the output of the absolute pressure sensor, which has not been determined to be in an abnormal state, the pressure in the second reference pressure chamber is calculated. The sensing system according to claim 1 or 2, which estimates the attitude state or motion state of the aircraft based on the output of the differential pressure sensor that has not been determined to be in an abnormal state and the pressure in the second reference pressure chamber.

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