Pressure sensor for determining airspeed, altitude and angle of attack
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to pressure measuring probes which provide multiple pressure readings from a single external strut mounted or boom mounted probe. 
     2. Prior Art 
     The advantage of having aerodynamically compensated static pressure sensing probes, as well as incorporating pitot sensing and dual static systems in the same probe has been known. Various devices which provide satisfactory results in many conditions have been advanced. For example, an aerodynamically compensated dual static probe is shown in U.S. Pat. No. 3,482,445. In that patent, the probe is provided with a discontinuity or surface irregularity that causes an alteration of the measured static pressure thereby compensating for other irregularities such as adjacent portions of the aircraft so that true dual compensated static sensing can be achieved by proper location of a pair of sets of ports on the probe. 
     In many applications it is necessary to measure local angle of attack on or ahead of the aircraft fuselage. In the present device, a second set of pressure ports are made on the side of the probe so that the pressure sensed will vary in a known relationship with respect to a first static pressure sensed at a first set of ports, and from this variation, the local angle of attack can be determined. 
     Both pitot-static tubes and angle of attack sensors have been used extensively in the past but generally each unit is installed at a different location thereby adding to the weight, drag and cost as well as the complexity of the installation. 
     Insofar as the location of the static pressure ports are concerned, U.S. Pat. No. 3,120,123 discloses a static pressure sensing probe which has ports located in various positions, including ports that are located at the sides of the probe and also an embodiment which shows static ports facing downwardly on opposite sides of the ventral central line. This configuration is shown in FIG. 15 of Pat. No. 3,120,123. 
     The ability to utilize differential pressure sensing ports on the hemispherical end of a sensing probe is disclosed in U.S. Pat. No. 3,318,146. The use of signals derived from such pressure sensing ports, by combining them in an air data computer, to provide angles of attack and sideslip is disclosed in that patent. 
     However, the present device permits the use of a slender probe for accomplishing the readings desired, which aids in minimizing drag, and also aids in simplifying the de-icing that is necessary. 
     SUMMARY OF THE INVENTION 
     The present invention relates to a sensing probe for use on aircraft, which may be small enough to meet design criteria from a physical standpoint, and yet can provide two separate sets of measuring ports that provide pressure measurements to obtain accurate static pressure and angle of attack readings. The ports are axially spaced and positioned in desired locations on the probe so that at a reference angle of attack they will be measuring the same static pressure or pressures at a known relation to each other. As the angle of attack either increases or decreases from the reference angle, the differential pressure sensed between the two port sets will change and this measured difference is used to determine the angle of attack. True static pressure can also be determined from these ports. 
     Additionally, the probe may be provided with a pitot pressure sensing port at its leading end in a known manner, to provide the pitot pressure reading directly with the same probe. 
     Each of the sets of pressure ports are located in a known relationship circumferentially around a separate diametral plane perpendicular to the axis of the probe, and the other of the sets of ports is spaced axially along the probe, and is not on the same diametral plane. However, by varying the pressure pattern along the probe as is now known, and positioning the static pressure ports in preselected locations, the second set of ports will be in a known relation to the primary static pressure sensing ports. 
     By using an angle of attack computer providing an output proportional to a known relationship between the differential pressure, static pressure and pitot pressure, the angle of attack can be provided directly. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a typical side elevational view of a fuselage mounted probe including ports located in accordance with the present invention; 
     FIG. 2 is a sectional view taken as on line 2--2 in FIG. 1; 
     FIG. 3 is a sectional view taken as on line 3--3 in FIG. 1; 
     FIG. 4 is a side elevational view of a typical boom mounted probe including static pressure ports located in accordance with the present invention; 
     FIG. 5 is a sectional view taken as on line 5--5 in FIG. 4; 
     FIG. 6 is a sectional view taken as on line 6--6 in FIG. 4; 
     FIG. 7 is a graphical representation of output pressures sensed by a probe of the present invention as a function of aircraft angle of attack; 
     FIG. 8 is a graphical representation of the pressure ratio between two sets of static sensing ports normalized in accordance with known procedures plotted versus the aircraft angle of attack and airspeed; and 
     FIG. 9 is a schematic representation showing the use of probes on opposite sides of an aircraft fuselage to eliminate effects of side slip. 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     In FIG. 1, a strut mounted probe illustrated generally at 10 is of the short probe variety, having a support strut 11, and a mounting base 12 that is generally fastened onto the fuselage 13 of an aircraft. The probe 10 includes a barrel section 14 that has a longitudinal axis, oriented in a preselected relationship with respect to the fuselage 13 of the aircraft. As shown, the barrel has a forward end 15, a first region indicated generally at 16 that is cylindrical in cross-section and of diameter D 1 , a second region indicated generally at 17 which includes a conical portion 18 joining the first region 16 and a cylindrical portion of diameter D 2  which is greater than D 1 . The forward end 15 is also tapered and at the leading end, a pitot pressure port indicated at 15A is provided for sensing pitot pressure. Further, as shown, an electrical heater element indicated generally at 21 may be mounted on the interior of the probe for de-icing purposes. 
     The interior of the barrel 14 is divided into individual sections, and suitable pressure carrying tubes lead from each of the sections. Bulkheads divide the probe to provide pressure sensing chambers on the interior of the probe. As shown in FIG. 2, a bulkhead 24 forms the leading end of the first static pressure sensing chamber. A first set of static pressure sensing ports 22, which as shown are on opposite sides and at equal angle θ 1  from the ventral longitudinal plane (on the lower side of the probe). The first ports are located in region 16 and provide pressure to a chamber of the probe. The pressure signal is carried through a suitable tube indicated generally at 23 to remote instruments. These instruments will be more fully explained subsequently. A pitot pressure sensing tube 25 extends through the bulkhead 24. Other bulkheads are used to provide a pressure sensing chamber for a set of ports 27 for example bulkhead 26 shown in FIG. 3. 
     The set of pressure sensing ports 27 is positioned in barrel section 18. As shown, ports 27 are diametrically opposed along a substantially horizontal plane. A line or tube 28 is used for carrying the second pressure sensed through the ports 27 to the remote instruments. Ports 27 are located in region 17. 
     As shown, the pressure sensed at ports 22,22 is normal static pressure provided to an altimeter 29 for altitude measurement. The pitot pressure supplied through line 25 is used in airspeed indication, and an airspeed indicator 32 is connected to this pitot pressure, and also to the line 23. 
     An angle of attack computer illustrated generally at 33, which is a ratio instrument giving an output along a line 34 to an angle of attack indicator 35, receives signals from the pitot pressure port 15A, and from both of the sets of pressure ports 22 and 27. The signals provided to the angle of attack computer may be electrical signals provided as functions of the measured pressures. Such electrical signals are provided as outputs of pressure sensors which deliver an electrical output as a function of pressure. These pressure sensors are well known and can form part of the angle of attack computer or they may be separate sensors used to provide electrical signals to other indicators or instruments. 
     The angle of attack computer is a ratio instrument well known in the art, and is programmed to provide outputs which are predetermined functions of the pressure measurements supplied to the computer, as will be further explained. 
     In relation to the device shown in FIG. 1, &#34;p tm  &#34; is the signal indicating the pitot pressure at port 15A, p m1  is the signal indicating the pressure at the first set of ports 22, and p m2  is the signal indicating the pressure at the second set of ports 27. 
     In FIG. 2, the angle θ 1  is generally about 371/2°. This arrangement provides a static pressure output which is substantially invariant with aircraft angle of attack. The pressure sensed at ports 27,27 comprising the second set of ports p m2  is equal to the pressure p m1  sensed at the set of ports 22,22 when the angle of attack is zero or at some other selected reference angle. However, the static pressure p m2  sensed at ports 27,27 changes with angle of attack. With the ports 27,27 positioned as shown, the pressure will decrease as angle of attack increases. 
     In dealing with sensed pressure quantities, a common term used is a normalized pressure function (p t  - p m )/q c . This normalized pressure function is the pitot pressure minus the measured static pressure divided by the impact pressure, q c . The quantity q c  is the pitot pressure minus the static pressure. The p m  designation means measured pressure and thus p m1  and p m2  are the measured pressures at port sets 22 and 27 respectively and p tm  is the measured pitot pressure at port 15A. These designations are used in the curves of FIG. 7. 
     Likewise, q cm1  as used in FIG. 7 is the measured pitot pressure minus the pressure measured at ports 22,22, that is q cm1  = p tm  - p m1 . In FIG. 7, the function using the pressure measured at ports 27, which is p m2  in FIG. 1, increases with increasing angle of attack and is illustrated by the curve 39. This plot of (p tm  - p m2 )/q cm1  shows a continuously increasing positive change with angle of attack of the aircraft when in the arrangement of FIG. 1. 
     FIG. 8 shows curve 40 which is a plot of (p m1  - p m2 )/q cm1  plotted versus indicated airspeed, and as a second plot, on the same figure, versus angle of attack. This ratio (p m1  - p m2 )/q cm1  is used by the aircraft angle of attack computer to provide an output proportional to angle of attack. In a critical region of the curve 40 shown in FIG. 8, that is in the region near stall, the sensitivity of the signal varies between 0.01 q cm1  and 0.04 q cm1  per degree, depending on the detailed design. Typically the pressures p tm , p m1  and p m2  are fed into two differential pressure sensors schematically represented at 45 and 46 in FIG. 1 which give electrical outputs proportional to the desired differentials p m1  - p m2  and p tm  - p m1 . 
     Referring now to FIG. 4, a nose mounted probe indicated generally at 50 is mounted onto the nose 51 of an aircraft, and as shown has a reduced barrel section 52, and a generally rearwardly increasing tapered forward section 53. A transition section 56 which ties the rear of the forward section 53 and the reduced section 52 together is also illustrated. 
     In this form of the invention, a pitot pressure port can be provided in the forward end of the probe and individual pressure sensing chambers for sensing two different pressures along the probe also are provided. The first, or primary static pressure sensing ports in this particular instance are rearwardly of the secondary sensing ports. As shown in FIG. 6, there are four (primary) static pressure sensing ports 54 located around the circumference of the probe, at equal angles θ 1  from the vertical bisecting plane. θ 1  in this case is generally equal to an angle in the range of 26°. 
     The secondary ports in this instance are located forwardly on the probe, and are shown at FIG. 5. Here, the ports 55 are at equal angles on opposite sides of the vertical bisecting plane and offset by an angle indicated at θ 2 . θ 2  in this instance is in the range of 10°, so that the ports 55,55 are more nearly directly at the bottom of the probe. 
     At a reference angle of attack, for example zero, the pressure sensed at each of the sets of pressure ports will be substantially the same or may be at a desired known relationship to each other. In a positive angle of attack, however, the set of ports 55,55 will sense an increased pressure in relation to the primary set of ports 54, and as shown in curve 58 in FIG. 7 the pressure function (p tm  - p m2 )/q cm1  (FIG. 4) will follow the trend shown in curve 58. That is, the pressure function will decrease with increasing angle of attack. For the plot of curve 58, q cm1  is p tm  (measured pitot pressure) minus static pressure measured at ports 54, and p m2  is the pressure measured at the second set of ports 55. 
     With the structure shown in the second form of the invention, the absolute values of the function |(p m1  - p m2 )/q cm1  | (p m2  would be the sensed pressure at the set of ports 55, and p m1  would be the sensed pressure at the set of ports 54) will be the same general shape as shown in FIG. 8 for curve 40, but the sign of p m1  - p m2  would be negative for the form of the invention shown in FIGS. 4, 5 and 6 since p m2  as shown would be larger then p m1 . 
     It should be noted that in the plot shown in FIG. 8, the output function is essentially linear and has a high slope in the region where angle of attack is most important, that is in the region of airspeed preceding stall. Commonly this linear region is from the stall speed to 1.4 times the stall speed for most aircraft. The system has therefore been provided with high sensitivity in the low speed region and lower sensitivity in the high speed region where angle of attack information is less important. 
     A further advantage of the system is that the pressures p m1  and p m2  can be manifolded from two identical probes 60 and 61 constructed as previously described with one on each of the right and left sides of the fuselage (FIG. 9) and the computed ratio is the average ratio or average angle of attack. This is important since for an aircraft fuselage at a combined angle of attack and sideslip, the local angle measured on either side of the fuselage will be altered by the sideslip. The local angle changes due to sideslip are nearly equal and of opposite sign such that the average reading from the two probes on opposite sides of the aircraft is the desired reading. 
     In a simplified system where a lift reserve signal would be sufficient (rather than angle of attack) the absolute value of the pressure differential (p m1  - p m2 ) could be transmitted to a direct reading instrument as described in U.S. Pat. No. 3,470,740. 
     It is apparent that either the leading or trailing sets of ports could be the ports at which the sensed pressure changes the most with changes in angle of attack.