Pressure Sensing Instrument for Aircraft

1. Field of the Invention

The present invention relates to an improved sensing instrument which utilizes pressure measurements during aircraft flight to determine the Mach number, flight angles, impact pressure, and static pressure with only one head protruding from the aircraft.

2. Prior Art

A pressure sensing instrument for aircraft is disclosed in U.S. Pat. No. 3,318,146, held by the same assignee. A device made according to the present invention is an improvement on the sensor disclosed in such patent. The device made according to the present invention exhibits marked improvements including an improvement in linearity wherein the output signal of the sensor is shown to agree very closely with the true angle of attack. A second significant improvement is that the present device is less sensitive to changes in Mach number of flow past the sensor. This is particularly true under conditions of subsonic flow.

FIG. 1 is a plan view of a sensor made according to the present invention and designed for mounting onto the fuselage of an aircraft.

FIG. 2 is a fragmentary enlarged end view of the end portion of the sensor taken as on line 2--2 in FIG. 1.

FIG. 3 is a fragmentary enlarged side view of the end portion of the sensor and adjacent portion of the sensor body taken as on Line 3--3 in FIG. 2 and including the addition of the prior art device shown by dashed lines.

FIG. 4 is a drawing of a spherical segment of one base.

FIG. 5 is a graph of nonlinearity flow angle error versus flow angle relative to the sensor in degrees comparing the '146 device and a device made according to the present invention.

FIG. 6 is a graph of sensitivity coefficient versus Mach number comparing the '146 device and a device made according to the present invention.

Referring to FIGS. 1 and 2 and the numerals of reference thereon, a sensor illustrated generally at 10 includes a probe 11 which is attached to an aircraft 12 and protrudes out into the air stream. The probe 11 has a suitable first support section 13 and a sensor body 14 comprising a cylindrical tube having an external surface about a central longitudinal axis 17, which longitudinal axis 17 is aligned with the axis of the aircraft in a predetermined relation.

End portion 15 of sensor body 14 comprises a part spherical surface defining a spherical section of one base having five ports thereon. As shown, leading end portion 15 has an axial port 16 in the center leading edge position on the extension of the central longitudinal axis 17 of sensor body 14. Longitudinal axis 17 is shown in FIGS. 1 and 3. End portion 15 has four other ports which are equally spaced in surrounding position with respect to axial port 16. Included are a first port 20 and a second port 21, both of which have their axes lying in a common plane, which plane passes through the longitudinal axis 17 of sensor body 14 and is vertical during straight and level flight. In addition, the axes of first and second ports 20 and 21 are preferably inclined at substantially 45 degrees from the longitudinal axis 17 of sensor body 14. The location of ports 20 and 21 has been found to be suitable between 35 and 55 degrees with respect to the longitudinal axis 17 of sensor body 14.

In addition, a third port 22 and a fourth port 23 are provided on end portion 15 and have their axes lying in a plane, which plane passes through the longitudinal axis 17 of sensor body 14 and perpendicular to the plane of the axes of ports 20 and 21. During straight and level flight, the sensor is oriented on aircraft such that the axes of ports 22 and 23 lie in the horizontal plane indicated at 25. The axes of ports 22 and 23 are also inclined at 45 degrees from the longitudinal axis of the sensor body 14 and these ports could likewise be positioned anywhere between 35 and 55 degrees with respect to the latter.

As shown, each of the ports are connected to a separate conduit. For example, port 20 is connected to a conduit 30, port 21 is connected to a conduit 31, port 22 is connected to a conduit 32, port 23 is connected to a conduit 33, and port 16 is connected to a conduit 34. These conduits open to their respective ports and extend through the sensor body 14 and the support section 13. The conduits can be supported on bulkheads or in any other suitable manner.

Assuming straight and level flight, it can be seen that the pressure at port 16 will be the total pressure, equal to the normal Pitot or stagnation pressure of an aircraft. The pressures at ports 20, 21, 22 and 23 will all be equal and more negative than the pressure at port 16 since they are situated downstream from the stagnation point. Now, if the angle of attack changes, a pressure differential occurs between ports 20 and 21, and the magnitude of this differential depends upon the amount of change in the angle of attack. Whether the differential is positive or negative would, of course, depend upon whether the change in the angle of attack is positive or negative. Likewise, if there is any sideslip, a pressure differential occurs between ports 22 and 23, and the measure of this differential is used in determining the angle of sideslip.

FIG. 3 illustrates the difference between a device made according to the instant invention and the prior art device. FIG. 3 is a side view of a portion of the device shown in FIG. 2 as taken from the right side of FIG. 2. Numbers in FIG. 3 identify the same components of the device as corresponding numbers in FIG. 2. The prior art device is shown by dashed lines superimposed on the present device. End portion 15 comprises a spherical segment of one base. Point of transition 48 between end portion 15 and sensor body 14 defines a circle lying in a plane perpendicular to the longitudinal axis 17 of sensor body 14. Center point 45 of the spherical segment defining end portion 15 is rearward of the plane of the point of transition 48. The end portion of the prior art device comprises a full hemisphere as shown by dashed lines 36. In the prior art device, the radius of the sensor body was equal to the radius of the end portion. The prior art sensor body is shown by dashed lines 38. Since the two radii were equal, there was a smooth transition at the point 49 where the end portion joined the sensor body, the sensor body being tangential to the end portion at the point 49 of joining.

As indicated above, end portion 15 of a device made according to the present invention comprises a spherical segment of one base. Such segment 40 is a part spherical surface and is depicted in FIG. 4. Segment 40 is the lesser of the two portions of sphere 42 that results from passing a plane through small circle 44 inscribed on the surface of sphere 42. As defined herein, the length of dimension 41, noted also as dimension "a", is in all embodiments less than the radius 43 of sphere 42. It follows that twice dimension 41 is less than the diameter of sphere 42 and that the surface of segment 40 is less than the surface of a hemisphere of sphere 42. The surface area of segment 40 is determined by the formulas S=2.pi.Rh=.pi.p.sup.2, where S equals surface area, R equals the radius of sphere 42 and the dimensions h and p are as shown in FIG. 4.

In the present application, dimension 41 equals the radius of sensor body 14, shown also as dimension 41 in FIG. 3. Since dimension 41 is less than the radius of sphere 42, shown as dimension 43 in FIG. 3, the point of transition 48 between end portion 15 and sensor body 14 is a locus of points defining a circle and transition 48 is sharp, unlike the smooth transition of the prior art device, shown at point 49. It is the sharp transition that results in the improved performance of a device made according to the present invention. It has been shown that such improved performance results where the radius of sensor body 14 is between 0.02 percent and 20 percent less than the radius of end portion 15.

The improvement in performance is graphically illustrated in FIGS. 5 and 6. The theoretically perfect instrument is one that is absolutely linear or in other words, the measured parameter is identical to the true parameter. While such theoretical linearity is rarely achieved, how closely an instrument approaches linearity is a measurement of its utility. If the instrument's deviation from linearity is within the error tolerance for a specific application, the instrument may then be utilized without corrections. Where the instrument's non-linearity is greater than that permitted by the error tolerance, the closer that the instrument is to being linear, the simpler are the corrections that must be applied to permit utilization of the instrument in the particular application. The theoretically linear instrument is shown in FIG. 5 as the true flow angle, line 50. The measured flow angles of an instrument made according to the present invention are shown by curve 52 and the measured flow angles of the prior art device are shown by dashed curve 54.

It can be seen that the present device, curve 52, closely approximates the true flow angle, line 50. The present device is dramatically better than the prior art device, curve 54, at the higher angles of attack. Modern military fixed wing aircraft frequently operate at high angles of attack in order to enhance their ability to quickly change direction. Such maneuvers may be critical to survival in a combat environment yet, such maneuvers are not without risk. The high angle of attack maneuvers place the aircraft close to the point of stall. Stall in a modern high performance jet aircraft is always critical since the aerodynamics of the aircraft may not permit recovery from a spin induced by the stall. Accordingly, a device made according to the present invention provides the pilot a very accurate indication, of angle of attack especially at the critical high angles of attack, thereby significantly contributing to flight safety.

FIG. 6 illustrates the Mach effects on both a device made according to the present invention, shown by curve 56, and a prior art device, shown by dashed curve 58. In FIG. 6 the theoretically perfect device is a sensitivity coefficient that is linear and does not vary as a function of Mach number. In the past, the greatest nonlinearity has occurred in the transonic region approaching Mach 1.0, with somewhat lesser nonlinearities occurring as Mach 1.0 is exceeded. Such nonlinearity, which frequently affects air data sensors, results from compressibility effects and the onset of shock wave formation as Mach 1.0 is approached. A device made according to the present invention is seen to be highly linear throughout a wide range of Mach numbers, including the critical transonic region, where only a slight nonlinearity exists. FIG. 6 shows the sensitivy coefficient, K, plotted against Mach number. The sensitivity coefficient is really a calibration constant used in the conversion of the various pressures sensed at the device to angle of attack, .varies., expressed in degrees. The relationship is normally expressed as ##EQU1## the various pressures, P, are the pressures sensed at the port shown in FIG. 2 with same number as the suffix of the specific pressure. That is, for example, P.sub.21 is the pressure sensed at second port 21. The near linearity of the present device can be seen from curve 56. It should be noted that in excess of Mach 1.0, curve 56 and curve 58 are coincident.

With this sensor we have five pressure measurements. The pressure at port 16 of P.sub.T ; and the pressures at ports 20, 21, 22 and 23. From these measurements the angle of attack, the angle of sideslip, the total pressure, the Mach number and the static pressure can be derived.

It has been found through analysis and experimentation that the Mach number of the airplane (Mach number is the ratio of the speed of the aircraft to the speed of sound in the fluid medium) is a function of the average pressure between the ports 22 and 23 (P.sub.4) divided by the stagnation pressure at port 16 (P.sub.T). It has also been found experimentally that this holds true across a wide range of Mach values. The derived curves shown in FIG. 6 represent substantially straight line functions, but vary in slope between three ranges of Mach values, namely, from 0.2 to about 0.8, from 0.8 to 1.1, and from 1.1 and over. Therefore, knowing the pressure ratio of P.sub.4 /P.sub.T, means is provided for determining the Mach number of the aircraft.

In addition, it has been found experimentally that the difference in pressure between ports 20 and 21 (P.sub.1 -P.sub.2) divided by the difference in pressure between port 16 (P.sub.T) and the average pressure at ports 22 and 23 (P.sub.T -P.sub.4) is a substantially linear function of the flow angle. Likewise, if the angle of sideslip is desired it can be determined by the ratio of the pressure differential between ports 22 and 23 (P.sub.1 and P.sub.2) divided by the difference in pressure between port 16 (P.sub.T) and the average pressure at ports 20 and 21 (P.sub.4 '). Accordingly, means is provided for determining both the angle of attack and the angle of sideslip.