Device and method for measuring the speed of a vehicle, in particular an aircraft

Abstract

A device is used to measure the speed of a vehicle, in particular the airspeed of an aircraft. The device has a sensor (10). This sensor has a pressure sensing element (20) with at least one opening (23). The pressure sensing element (20) has this opening (23) on a side facing away from the direction of movement, in particular the direction of flight, of the aircraft. The pressure sensing element (20) with the opening (23) is designed and arranged such that a measurement of the negative dynamic pressure is taken.

Description

[0001] The invention relates to a device and a method for measuring the speed of a vehicle, in particular an aircraft, using a sensor.

[0002] The speed of aircraft is important for many questions. This includes not only determining the time an aircraft needs to travel from point A to point B, but also the speed at which an aircraft moves relative to the surrounding air mass. This is crucial, for example, for determining how the aircraft's rudder and wings should be moved or adjusted to optimally reach a specific destination or even to create the conditions necessary for the aircraft to move in the first place.

[0003] Knowing this speed relative to the air, also known as airspeed, allows us to determine further physical values ​​using additional data such as wind speed. However, the basis for this calculation is the aforementioned uncorrected airspeed of the aircraft, also known as IAS (Incidence Airspeed).

[0004] A pitot tube, also known as a Prandtl tube, is typically used to measure this speed. This pitot tube measures the pressure difference between the static pressure of the air flowing around the aircraft and the so-called total pressure, which is the sum of the static and dynamic pressures. This measurement method provides a good basis for assessing airspeed and has therefore been used for decades.

[0005] For example, EP 2 434 296 B1 describes a special airspeed sensor system for an aircraft with several pitot tubes.

[0006] A prerequisite for a functioning airspeed measurement using a pitot tube is that the pitot tube is unobstructed, unaltered, and fully functional. In practice, foreign objects such as bird feathers or insects, but especially snowflakes, hailstones, or large water droplets, can clog the pitot tube. Functionality is regularly impaired when, under certain external conditions, large water droplets or other foreign objects in the surrounding air mass enter the pitot tube, and particularly when icing occurs. This can lead to significant malfunctions in the entire system. If the pitot tube is narrowed or completely blocked, the precision of the pressure measurement is no longer guaranteed.Icing in the clogged pressure measuring system leads to significant false readings and incorrect or even completely wrong initial values ​​for further calculations of derived quantities and control variables in connected controllers.

[0007] In fact, this problem has already led to accidents in air traffic practice.

[0008] This problem of pressure-based airspeed measurement in aircraft, which is prone to icing, is also being addressed. For example, the pressure gauge tube, or pitot tube, used to measure dynamic pressure is heated for this purpose. This heat input evaporates any water particles entering the pitot tube, thus reliably de-icing icy surfaces. The additional weight of the necessary heating equipment does not pose a significant problem for larger commercial aircraft. Furthermore, the energy required for the heat input is drawn from the energy source used to power the aircraft's engines and other electrical components on board, such as the fuel.

[0009] Although the use of appropriate heating devices in pitot tubes for measuring aircraft airspeed has been widespread and for a long time, the question remains whether there is another way to solve the problem of icing-prone pressure-based airspeed measurement on an aircraft. The object of the invention is therefore to propose a generic device and a method that solve the problem in an alternative way.

[0010] This problem is solved in a generic device for measuring the airspeed of aircraft by the fact that the sensor has a pressure sensing element with at least one opening, that the pressure sensing element has the opening on a side that is facing away from the direction of movement of the vehicle, in particular the aircraft, and that the pressure sensing element with the opening is constructed and arranged in such a way that a measurement of the negative dynamic pressure is carried out.

[0011] In one method, this problem is solved by taking a pressure reading from a pressure sensing element at an opening, on a side that is opposite to the direction of travel of the vehicle, in particular the aircraft, and by measuring the negative dynamic pressure.

[0012] The pressure measurement for determining the flight speed is then taken at a location facing away from the direction of flight and thus away from the direction of incidence of water droplets and other foreign bodies.

[0013] This design surprisingly provides a solution. While prior art inevitably accepts that a problem will arise from a blockage and then be resolved, or at least requires continuous and necessary work to combat any emerging problem, the invention specifically prevents the problem from occurring in the first place. This means that a heating system is no longer required. Consequently, the heating system is eliminated as an accessory, indeed, as an entire component. The energy consumption of such a heating system no longer needs to be considered or supplied, nor does it require regular maintenance and monitoring of its function. Such maintenance and monitoring are quite labor-intensive in conventional aircraft, as malfunctions must be prevented.

[0014] While the costs and weight of the heating system may be manageable, they do lead to significant increases in the load, especially in the case of so-called drones or small aircraft.

[0015] According to the invention, the elimination of the heating element is replaced by a simple device for measuring or recording the pressure. Such an additional device does not result in any significant increase in weight or cost.

[0016] Such a device according to the invention for measuring the airspeed of aircraft is particularly suitable for lighter and smaller aircraft. Especially in lighter and smaller aircraft, energy consumption and the weight of a heating system are of particular importance, for example, in the case of drones, also known as uncrewed aircraft systems or unmanned aircraft systems (UAS). In these cases, every energy saving is of particular value, especially since no maintenance or repair of the heating system is required.

[0017] The use of heating or the provision of a heating device is not precluded by the inventive design. There may be applications where such an additional heating device is advantageous. For example, vertical take-off and landing (VTOL) aircraft can be considered. In these aircraft, the pressure sensors are sometimes exposed to airflow from the side normally facing away from the direction of flight.

[0018] In unmanned aviation with the aforementioned drones, a further advantage arises, particularly for a concept according to the invention. In unmanned aviation, the acceptable error tolerances in airspeed measurement are typically greater compared to the requirements in manned aviation.

[0019] This means that, due to the design, more cost-effective and somewhat less precise pressure measurements can be used, since the accuracy of a classic heated pitot tube is not absolutely necessary to meet the requirements. What is relevant is achieving a sufficiently precise result within acceptable error tolerances for the airspeed measurement, and a somewhat less expensive version of the device according to the invention also satisfies these requirements.

[0020] For the functioning of the device according to the invention for measuring the airspeed of aircraft, the possibility of using the measurement of the dynamic pressure, which can correspond to a negative dynamic pressure at specific points around a body in a flow, instead of the positive ram pressure at a pitot tube from the prior art is relevant.

[0021] Negative ram pressure, in other words the dynamic pressure at negative pressure coefficients or negative dynamic pressure, which thus uses the suction effect, is unexpected at this point, as it has not yet been used to measure the forward speed of an aircraft and has not been considered.

[0022] Negative dynamic pressure is already known for entirely different purposes in aviation. For example, it is used in total energy compensation in gliding, where it is achieved, for instance, with a TEK nozzle. Special designs of such nozzles are also known as pitch or twin-slot nozzles. However, they have not yet been used for robust airspeed measurement under icing conditions, and their application is surprising.

[0023] According to the invention, the pressure coefficient of the probe used becomes an important parameter.

[0024] The following applies: pdyn=pstau×cpsonde where p dyn the dynamic pressure, p stau the dynamic pressure and c psonde The pressure coefficient of the probe is important. The pressure coefficient should be as constant as possible. One embodiment with c is particularly easy to understand. psonde = -1, but other values ​​are also possible, apart from c psonde = 0.

[0025] With state-of-the-art heated probes, high energy consumption is a significant disadvantage. The heating element can malfunction or fail completely. At low temperatures and / or high flight speeds, the heating performance can be unsatisfactory. Furthermore, a forward-facing pitot tube can become blocked not only by precipitation, rain, or ice, but also temporarily by dirt, leading to inaccurate readings.

[0026] In contrast, the invention enables a robust and extremely simple pressure measurement for determining airspeed, even under adverse conditions. This includes, for example, Saharan dust, volcanic ash, rain and snow, and of course the icing conditions mentioned above.

[0027] A particular application of the device according to the invention for measuring the airspeed of aircraft is its use with drones, for example, for measuring vertical profiles of the entire troposphere. In such cases, very extreme weather conditions must be taken into account, and the use of the measured vertical profiles is particularly undesirable under such conditions, since the determination of values ​​is important not only for the current flight but also for further meteorological applications, for example, for national weather services.

[0028] The idea can be implemented particularly practically with a method in which the static pressure is also measured, and in which a differential pressure is determined from the values ​​of the negative dynamic pressure and the static pressure by direct mechanical measurement of the differential pressure or by determination from the measured values ​​of the individual pressures.

[0029] A preferred probe design is one in which the pressure sensing element is shaped like a spherical probe and the opening on the spherical surface is positioned at an angle between 90 and 170 degrees to the flight direction of the aircraft or other vehicle. However, the opening is not directly opposite to the commonly used pitot tube, but rather angularly offset. Angles between 110 and 120 degrees, aligned with the flight direction of the vehicle, are particularly preferred.

[0030] The concept according to the invention makes it possible to use the system even in aircraft under environmental conditions where icing would conventionally occur. This can also be achieved by comparing the use of a conventional, but unheated, pitot tube alongside the concept based on the invention, which actively detects and prevents icing conditions.

[0031] One interesting area of ​​application is, for example, weather drones for national and other weather services and also for other meteorological applications.

[0032] Various areas of an aircraft are suitable locations for the placement of the sensors as part of the device for measuring airspeed. For example, a placement in the center of the aircraft's nose is conceivable, or more generally, at the front of the aircraft.

[0033] Another option would be a design with a kind of bulge on the hull.

[0034] An arrangement at the rear, for example again in the middle, is also possible.

[0035] However, other arrangements are also conceivable.

[0036] It is particularly advantageous if the sensor has a differential pressure sensor, if the further opening is arranged in such a way that it absorbs the static pressure, and if the differential pressure sensor is designed in such a way that it compares the two pressure values.

[0037] This concept is particularly suitable for implementation using a method in which a measurement or determination of, on the one hand, a positive dynamic pressure and, on the other hand, a differential pressure between negative dynamic pressure and stationary pressure is carried out, in which these values ​​are compared, and in which a conclusion is drawn from this comparison regarding the icing state.

[0038] Further preferred features and advantages of the invention are explained in more detail with reference to the patent claims and the following exemplary embodiments.

[0039] The invention is intended and applicable primarily for aircraft moving in a medium typically formed by the Earth's atmosphere. However, in principle, the concept according to the invention is also conceivable for vehicles moving in other media, including underwater or extraterrestrially on other planets. Therefore, in addition to air, other fluids that cause similar problems, which are solved by the invention, can also be considered.

[0040] An embodiment of the invention is described in more detail below with reference to the drawings.

[0041] They show: Fig. 1 a section through a schematic representation of an embodiment of the invention; Fig. 2 a representation of the arrangement of an embodiment of the invention on an aircraft; Fig. 3 a schematic representation of an arrangement of an embodiment of the invention, which creates further possibilities; and Fig. 4 a schematic representation similar to the Fig. 3.

[0042] In the Fig. Figure 1 illustrates a sectional view of a sensor 10 as used in a device for measuring the airspeed of aircraft.

[0043] The illustration is purely schematic and is intended to explain the principles used according to the invention. In the illustrated embodiment, the sensor 10 has the form of a probe and includes a differential pressure sensor 11, which is preferably arranged and attached directly to the outer skin of the aircraft.

[0044] Inside this differential pressure sensor 11, two cavities 12 and 13 are provided. These are located in the Fig. The diagrams are shown side-by-side, but other configurations are also possible. The two cavities 12 and 13 are arranged and connected in such a way that the differential pressure between the two cavities 12 and 13 can be measured and the data thus obtained can be further processed via a transducer 14.

[0045] Both cavities 12 and 13 are connected to the aircraft's environment. This is shown schematically. For this purpose, the differential pressure sensor 11 is equipped with a pressure sensing element 20. The pressure sensing element 20 extends from the differential pressure sensor 11 parallel to the outer skin (not shown in the diagram). Fig. 1) and then bent 90 degrees outwards into the surrounding area away from the aircraft.

[0046] The pressure squeezing element 20 is located near its outermost end, facing away from the aircraft and the two cavities 12 and 13, and is provided with an opening 23 that opens in the opposite direction of flight. In its inner region, the pressure squeezing element 20 is provided with a cavity 21, which is surrounded by a jacket 22, such that the opening 23 in the jacket 22 connects the interior, or cavity 21, with the environment of the aircraft. Since the opening 23 is located on the side of the tubular pressure squeezing element 20 facing away from the ambient gas flowing around the aircraft, pressure conditions corresponding to negative dynamic pressure are established in the interior, or cavity 21.

[0047] In a section of the pressure sensing element 20 between the interior 21 and the differential pressure sensor 11 with the two cavities 12 and 13, two further openings 33 are provided. These two further openings 33 are not connected to the interior 21. The interior 21 is instead connected to the [unclear text] via a thinner line 24. Fig. 1. The cavity 12 shown on the left is connected. The pressure conditions in the cavity 12 therefore correspond to the negative dynamic pressure at the opening 23 for pressure relief.

[0048] The openings 33 of the pressure sensing element 11 in the area between the interior 21 and the cavity 12 also lead to lines, in this case to two schematically depicted lines 34, which transmit the existing pressure conditions to the right in the Fig. Pass on the cavity 13 shown in section 1.

[0049] The openings 33 in this area of ​​the pressure sensing element are not subjected to dynamic pressure, but only to static pressure.

[0050] The following relationships should be noted: v = flow velocity ϱ = Air density p S = static pressure p E = Replacement pressure pE=ps+pdyn pdyn=ϱv2 / 2×cpsonde p dyn = dynamic pressure c psonde = Pressure coefficient at the measuring point, here x - 1 (dimensionless)

[0051] The differential pressure sensor 11 can also be placed via a different arrangement of lines 24 or 34, in particular also via hose lines.

[0052] In the Fig. Figure 2 shows a schematic representation of the arrangement of an embodiment of the invention on an aircraft. Three different embodiments of a device are indicated here.

[0053] The image shows an aircraft, depicted here purely schematically for illustrative purposes; for example, a passenger plane. Other possibilities include drones or other aircraft.

[0054] The image shows a fuselage 41 with windows 42. Wings 43 can be seen in the section of the fuselage.

[0055] Below the hull 41, one can see the outlines of bow wheels 44 and a stern wheel 45.

[0056] It is specifically indicated that a sensor 10 may be located at the foremost end of the fuselage 41, i.e., at the nose of the aircraft, as detailed in the Fig. 1 is shown.

[0057] Instead of or in addition to the sensor 10 shown, a sensor 10 may also be located at the rear of the aircraft or in the middle area of ​​the fuselage, in particular in the middle of the aircraft 40 and in the plane of symmetry of the aircraft 40.

[0058] It can be seen from the indicated embodiments of the sensors 10 that the openings 23 in the casing 22 of the pressure sensing element 20 of the sensors 10 are located on the side of the sensors 10 facing away from the wind. This ensures that negative dynamic pressure prevails at this point and can be utilized.

[0059] In the Fig. Figure 3 presents a further development of the invention which makes use of the inventive ideas.

[0060] Here, pressure reduction occurs via a pressure reduction element 20 in the form of a spherical probe. The spherical probe of the pressure reduction element 20 is oriented opposite to the flow direction of the surrounding atmosphere. In practice, it is subjected to supercritical flow. The Reynolds number is greater than 30,000.

[0061] The spherical probe of the pressure sensing element 20 has openings 23 located at approximately 110° to 120°. These are pressure bores. The positioning of the openings 23 can also be at other angles. The angles are relative to a value of 0°, which would be chosen for an opening exactly in the direction of flow.

[0062] The ball probe of the pressure sensing element 20 thus possesses, like the concept of the pressure sensing element 20 from the Fig. 1 Openings 23 on a side facing away from the direction of flight of the aircraft. In this case as well, the arrangement of the openings 23 ensures that no interfering particles can penetrate into the openings 23 (in this case in the spherical probe of the pressure sensing element 20), i.e. no bird feathers or insects and in particular no water droplets that cause snowflakes or icing.

[0063] Openings 33 for measuring static pressure are also provided in the casing 32. Here again, a combination of lines is provided to enable robust pressure measurement.

[0064] This is p = p dyn × c psonde c psonde approximately -0.7. This value needs to be calibrated.

[0065] The elements described above already constitute an interesting embodiment for a concept with a specially designed pressure sensing element 20. The differential pressure sensor 11 shown and described from the embodiment in the Fig. 1 can be formed in this embodiment by an electronic unit to which the respective measured values ​​are supplied. Of course, a combination in another form is also possible, i.e., an electronic unit for comparing the measured values ​​in the embodiment of Fig. 1 and a more mechanically conceived design of the differential pressure sensor 11 in a representation otherwise shown in Fig. 3 corresponding embodiment.

[0066] The pressure sensor can therefore be connected in such a way that either the differential pressure of the two pressure values ​​is measured, or individual pressures are measured, or both differential and individual pressures are measured by distributing the pressure lines. As a result of the various configurations, not all of which are shown in the figures, the differential pressure is either measured or calculated electronically from the measured values.

[0067] The following describes a further embodiment that utilizes this concept with the pressure sensing element 20 in the form of a spherical probe. An embodiment is also conceivable, though not shown, in which one individual value is determined with such a spherical probe, while another individual value is determined with a pressure sensing element 20 designed as shown in the Fig. 1 is combined.

[0068] However, the following will show the representation in the Fig. 3 described.

[0069] Additionally, another interesting element is provided here, namely a conventional Pitot tube 50. A pressure sensing element is deliberately provided here, which has its opening opposite the direction of flow.

[0070] This Pitot tube is therefore affected precisely by the icing conditions which, incidentally, are avoided according to the invention.

[0071] This pressure sensing element, or pitot tube 50, now determines the pressure conditions in the conventional manner using a pressure sensing element 51. The following applies approximately: p dyn or p total as a result of the pressure decrease.

[0072] This cleverly allows for a comparison of a conventional pressure measurement, affected by icing, with all its consequences, against a pressure measurement according to the invention, which is not affected by icing, and the resulting velocity values. By comparing the respective data, it can be reliably determined whether they correspond or whether a deviation actually occurs. If a deviation is detected, this can only be due to icing or similar defects. This means that an immediate signal and corresponding data indicating the presence of icing are obtained. From this, conclusions can then be drawn about other effects of this icing on the aircraft. Thus, icing detection is combined with a highly precise velocity measurement.This can be provided in one and the same device, thus requiring relatively little effort.

[0073] It is also conceivable to use two Pitot tubes 50 with different susceptibility to icing (not shown), which would therefore have different susceptibility to icing. For safety reasons or other considerations, one might also perform further pressure measurements of this type with varying degrees of icing susceptibility and then compare the resulting values ​​to gain additional insights, for example, into the type of icing. Pitot tubes with different opening diameters, for instance, have different susceptibility to icing.

[0074] The comparison between devices that react differently to icing due to a decrease in pressure thus makes it possible to determine icing conditions.

[0075] Furthermore, other embodiments are conceivable, in which a comparison device for cabin pressure is also provided. Conventional pressure measurements in manned light aircraft sometimes switch to cabin pressure when the static pressure port ices up, since icing is largely ruled out in this case and approximate measurements seem possible. A comparison using these values ​​may therefore be useful or desirable for specific applications.

[0076] An alternative embodiment of an embodiment of an arrangement according to the invention is described in the Fig. 4 shown.

[0077] Instead of direct absolute pressure measurement or differential pressure measurement, the differential pressure can be determined using a flow sensor 62. This utilizes the relationship between pressure loss and flow rate that arises when fluid passes through a throttle 61 or capillary. In combination with the pressure drop at a location where the dynamic pressure is negative, the air (or fluid in general) flows out, while the fluid flows in at the point of pressure drop for the static pressure. Under extreme icing conditions, especially when flowing along the body, the outflowing air at the point of pressure drop for the dynamic pressure can further increase robustness, provided the point of pressure drop for the static pressure is adequately protected. If the static pressure is taken, for example, inside an aircraft, the robustness of the measurement can be further increased by the outflow of warm air at the sensor head.The flow rate can be measured, for example, via calorimetric flow measurement (measurement via heat transport in the fluid).

[0078] One can see in the Fig. 4 in turn the openings 23 for dynamic pressure and 33 for static pressure.

[0079] In addition, a throttle 61 or capillary and a calorimetric flow sensor 62 are included. This again results in a differential pressure measurement, in this case via the throttle 61, and a flow measurement with the flow sensor 62. For better understanding, the circuit symbol for a volumetric flow meter 63 is also shown next to the flow sensor 62. This also creates a comparator device 52 in this case. Reference symbol list 10 Sensor 11 Differential pressure sensor 12 Cavity in the differential pressure sensor 13 Cavity in the differential pressure sensor 14 transducers 20 Pressure squeezing element 21 Interior of the pressure squeezing element 22. Sheath of the pressure sensing element 23 Opening 31 Management 32 coat 33 Opening for static pressure 40 aircraft 41 Fuselage of the aircraft 42 windows 43 wings 44 front wheels 45 rear wheel 50 Pitot tube 51 Pressure sensing element 52 Comparator device 61 Throttle 62 Flow sensor 63 Volume flow meter QUOTES INCLUDED IN THE DESCRIPTION

[0000] This list of documents cited by the applicant was automatically generated and is included solely for the reader's convenience. The list is not part of the German patent or utility model application. The DPMA accepts no liability for any errors or omissions. Cited patent literature

[0000] EP 2 434 296 B1

[0005]

Claims

[1] Device for measuring the speed of movement of a vehicle, in particular an aircraft, within a fluid, in particular within the Earth's atmosphere, comprising a sensor (10), characterized by , that the sensor (10) has a pressure sensing element (20) with at least one opening (23), that the pressure squeezing element (20) has the opening (23) on a side which is turned away from the direction of travel of the vehicle, in particular the aircraft, and that the pressure sensing element (20) with the opening (23) is constructed and arranged in such a way that a measurement of the negative dynamic pressure is carried out. [2] Device according to claim 1, characterized by , that the sensor (10) has a differential pressure sensor (11), that the further opening (33) is arranged in such a way that it absorbs the static pressure, and that the differential pressure sensor (11) is designed to compare the two pressure values. [3] Device according to claim 2, characterized by , that the differential pressure sensor (11) is of a mechanical or electronic design, and that the measured values ​​of the pressure sensing elements (20) are supplied to him as individual values ​​and / or as a total measured value. [4] Device according to any one of the preceding claims, characterized by , that the pressure sensing element (20) has the shape of a spherical probe and that the opening (23) on the spherical surface of the spherical probe is arranged in a position which encloses an angle between 90 degrees and 170 degrees, in particular between 110 degrees and 120 degrees, with the direction of flight of the aircraft. [5] Device according to any one of the preceding claims, characterized by, that an additional pressure sensing element (51) is provided, and that a comparator device (52) is provided which is designed to receive the pressure values ​​of the pressure sensing elements (20) and (51). [6] Device according to claim 5, characterized by , that several pressure squeezing elements (20) and (51) of different susceptibility to icing are provided, and that the comparator device (52) compares the data of all pressure sensing elements (20) and (51). [7] Method for measuring the speed of movement of a vehicle, in particular the airspeed of an aircraft, within a fluid, in particular within the Earth's atmosphere, characterized by , that pressure is taken off at a pressure sensing element (20) at an opening (23), on a side that is opposite to the direction of travel of the vehicle, in particular the direction of flight of the aircraft, and that a measurement of the negative dynamic pressure is performed. [8] Method according to claim 7, characterized by , that, in addition, the static pressure is measured, and that a differential pressure is determined from the values ​​of the negative dynamic pressure and the static pressure by direct mechanical measurement of the differential pressure or by determination from the measured values ​​of the individual pressures. [9] Method according to claim 7 or 8, characterized by , that a measurement or determination of, on the one hand, a positive dynamic pressure and, on the other hand, a differential pressure between negative dynamic pressure and stationary pressure is carried out, that a comparison of these values ​​is made, and that a conclusion is drawn from this comparison regarding the state of icing.

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