Wind-tunnel investigation of a flush airdata system at Mach numbers from 0.7 to 1.4

NASA Technical Memorandum 101697

Wind-Tunnel Investigation of a Flush

Airdata System at Mach Numbers

From 0.7 to 1.4

ilil I

Terry J. Larson, Timothy R. Moes, and Paul M. Siemers III

II April 1990 t N90-18395 (NASA-TM-IO]6?7) WIND-TUNNEL INVESTIGATION OF A FLUSH AIROATA SYSTEM AT MACH NUMBERS FROM 0.7 TO 1.4 (NASA) 35 p CSCL OlD Unclas G3/06 0270349

NASA

National Aeronautics and Space Administration NASA Technical Memoradum 101697

Wind-Tunnel Investigation of a Flush

Airdata System at Mach Numbers

From 0.7 to 1.4

Terry J. Larson Analytical Mechanics Associates, Hampton, Virginia Timothy R. Moes Ames Research Center, Dryden Flight Research Facility, Edwards, California Paul M. Siemers III Langley Research Center, Hampton, Virginia w

I_I/LSA

National Aeronautics and Space Administration Ames Research Center Dryden Flight Research Facility Edwards, California 93523-0273 CONTENTS ABSTRACT INTRODUCTION NOMENCLATURE DESCRIPTION OF MODEL TEST INSTRUMENTATION AND CONDITIONS 3 RESULTS AND DISCUSSION CONCLUDING REMARKS 5 REFERENCES TABLES FIGURES 18 !11 PRECEDING PAGE BLANK NOT FILMED

ABSTRACT

Flush pressure orifices installed on the nose section of a l/7-scale model of the F-14 airplane were evaluated for use as a flush airdata system (FADS). Wind-tunnel tests were conducted in the 11- by 1l-ft Unitary Wind Tunnel at NASA Ames Research Center. A full-scale FADS of the same configuration was previously tested using an F-14 aircraft at the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden). These tests, which have been published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind-tunnel test MaCh numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2° increments from -4 ° to 20 ° . Sideslip angles were varied in 4 ° increments from -8 ° to 8 ° .

Airdata parameters were evaluated for determination of free-stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

INTRODUCTION

For the past several years the Dryden Flight Research Facility of NASA Ames Research Center (Ames-Dryden) has been conducting a wind-tunnel and flight test program to evaluate the accuracy of airdata systems comprised of only pressure orifices flush with the aircraft skin. Such a flush airdata system (FADS) is intended to provide all the conventional airdata quantities, including angle of attack and sideslip, generally provided by pitot-static and flow-angularity-measuring systems on aircraft.

The FADS investigation was begun as support for the shuttle entry airdata sensor (SEADS) program conducted by the NASA Langley Research Center (Siemers and Larson, 1979; Siemers et al., 1988). The SEADS is a FADS- type experimental sensor originally developed to provide airdata for the orbiter at hypersonic speeds. However, early in the program a wind-tunnel test indicated that SEADS could also be accurately applied at lower speeds CLarson and Schweikhard, 1978).

Flight and wind-tunnel data have been obtained at subsonic, transonic, and low supersonic speeds to provide a means for evaluating the use and accuracy of FADS for general flight applications. In addition, these data helped validate the SEADS concept and data algorithms for application at the conditions tested. The subsonic data were obtained using a KC-135 airplane and scale model CLarson et al., 1980; Larson and Siemers, 1980, 1981), and the higher speed data were obtained using an F-14 airplane CLarson et al., 1987). FADS data were also obtained on a 1/7-scale model of the F-14 airplane in the 11- by i l-ft unitary model at the Ames Research Center (Research Facilities Handbook, 1982).

The FADS for the F-14 wind-tunnel tests, as for the flight tests, consisted of a spherical nose-cap surface with 11 pressure orifices in a cruciform pattern and additional orifices on the fuselage nose. Wind-tunnel tests were made at Mach 0.60, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied at 2 ° increments from -4 ° to 20 °, and sideslip was varied in 4 ° increments from -8 ° to 8 °. The Mach number and flow-angle ranges for this test were not sufficient for complete calibration over the F- 14 flight envelope, but the ranges were sufficient to evaluate the FADS configuration.

The FADS data from the wind-tunnel test are cited in this report. All the data, except repeat runs, are presented in tabular form. Representative data are discussed for application to airdata determination. The criteria used in Larson ct al. (1987) for sclection of the airdata parameters were also used in the wind-tunnel investigation.

NOMENCLATURE

FADS flush airdata system free-stream Mach number M free-stream static pressure, lb/in 2 P, free-stream total pressure, lb/in 2 pt stagnation pressure, lbAn 2 Pt2 pressures measured at orifices 1 to 11 on nose cap, lb/in 2 pl to pll pressures measured at nose section orifices 12 to 23, lb/in 2 pl2 to p23 free-stream dynamic pressure, lb/in 2 SEADS shuttle entry airdata system O_ angle of attack, deg

# angle of sideslip, deg

FADS Pressure Parameters: angle-of-attack parameter, nose cap (p7 - p3)/_ APn AP12 angle-of-attack parameter, nose cap (p6 - to3)/?/ angle-of-attack parameter, nose cap (197 - Pl)/q APn APl4 angle-of-attack parameter, nose cap (P6 - lo2)/] angle-of-attack parameter, station 61 (pt4 - p12)/q AP:I AP3_ angle-of-attack parameter, station 127 (p2o - pl6 )/t7

BPtt angle-of-sideslip parameter, nose cap (Ps - pll)/q

BP12 angle-of-sideslip parameter, nose cap (p9 - plo)/q

B angle-of-sideslip parameter, station 61 (P15 - pl3)/q

BP32 angle-of-sideslip parameter, station 127 [ (P21 + lo23) - (pit + pig) ]/2 t7 Mach number parameter, pl/p4

MP1

MP5 Mach number parameter, 2 p4/(P13 + P15 )

Mach number parameter, 2p4/(Pls + p22)

MP6

PP22 static-pressure parameter, station 61 (pl2 + p14)f2P6 static-pressure parameter, station 127 (P16 + p20)/2p_ PP32 PP33 static-pressure parameter, station 127 (pl6 + pÂ7 ... + 1923)/8ps PTPI4 stagnation-pressure parameter, nose cap, pa/Pt PTP15 stagnation-pressure parameter, nose cap, ps/Pt stagnation-pressure parameter, nose cap, p6/pt PT P16

DESCRIPTION OF MODEL

The 1/7-scale metal model was equipped with inlets for performance testing. A photograph of the model installed in the wind tunnel is shown in figure 1. The model was considered adequate for testing the FADS concept. However, adirectcomparison ofthetestresults withfull-scale testdatawouldbeexpected tohave some significant differences because of configuration differences between theairplane andthemodel. Thefactthatthemodel is wingless should havesome effect.However, because theaircraft's wingsarewcll aftof thenosesection, theeffectmaybesmaller thanthose of otherconfigurational differences. Forexample, theairplane hadcanardlike surfaces onthefuselage forcbody forspinrecovery thatcouldhave affected theflightFADS measurements. Eventhough these surfaces were closed forthetests, theywerenotflushwiththeairframe. Unlikethcmodel, theaircraft included (1)agunfairingon theleft sideof thefuselage, (2)aninfrared scanner podunderneath theairplane's nose, and(3) atestairdata probe attached tothenose of thepod.Forfurther details oftheairplane, see Larson etal.,1987.

Aswith thetestairplane (Larson etal., 1987), theendof themodel's nose wasreplaced with a sphere-shaped nose capcontaining 11pressure orificesarranged in a cruciformpattern.Thegeometrical spacing andassigned numbering areshown in figure2.Thesensor is essentially a l/7-scale replica of theflighttestsensor exceptforthe orificescaling. Themodel orificediameters were0.0134 in,,whereas theairplane orifices were0.0311 in. A smaller model orificesizewasavoided because of possible particulate plugging of theorifices.

Tworowsofpressure orifices wereplaced around themodel onthefuselage nose aftof thenose capasshown in figure 2. These locations areatthesame relative locations asthetestairplane installations. Thefirstrowis referred toasbeingatstation 61andthesecond rowatstation127, whichweretheoriricedistances in in. aftoftheairplane's nose apex.Fourorificesatstation127ontheairplane werenotincluded onthemodel.These wereorifices24,25, 26,and27 (Larson etal., 1987), located onthebottom of thefuselage. Theflighttestsindicated thatthese closely spaced orificesdid not provide any additional information not already provided by the other orifices.

TEST INSTRUMENTATION AND CONDITIONS

The FADS wind-tunnel testing was performed in the NASA Ames Research Center's 11 - by 11-ft Unitary Wind Tunnel. The model was sting mounted as shown in figure 1. Tables 1 and 2 show the test conditions.

All 23 FADS pressures were measured by 0-15 lb/in 2 transducers having accuracies of 0.25 percent of full _cale. Pressures were recorded with a mechanically operated scanning system. The overall accuracy of the pres- sure measurements were within 0.045 lb/in 2. The tunnel testing characteristics are reported in Research Facilities Handbook, 1982.

All test data, with the exception of repeat runs, are presented in tables 1 and 2. The FADS pressures for the repeat runs generally varied less than 0.002 lb/in 2 . Table 1 presents the data for 0 ° angle of sideslip. Note that the first 5 quantities are the test conditions, while the remaining 23 quantities are the ratios of FADS pressure to total pressure. The first group of data are for the various Mach numbers near 0.73. The test Mach number for this group will be referred to as 0.73 even though Mach numbers varied during the angle-of-attack sweep. This group of data is followed by similar groups of data for test Mach numbers of 0.90, 1.05, 1.20, and 1.39. Again, it is noted that Mach numbers varied, in some cases as much as 0.06, from the test values.

Table 2 presents the test data for angles of sideslip of -8 °, -4 °, 0 °, 4°, and 8°. Note that in most cases three angles of attack were obtained for each sideslip angle.

RESULTS AND DISCUSSION

Airdata parameters are shown in figures 3 to 7. These five figures present parameters for determining stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number. All of these figures are for 0 ° angle of sideslip except for figure 6, which shows the variation of the sideslip parameters with angle of sideslip. The data for each figure are plotted against angle of attack or sideslip for each of the five test Mach numbers.

According to theanalysis, thethree most suitable stagnation-pressure parameters were PTPt4, PTP15, and PTP16, calculated using orifices 4, 5, and 6, respectively. These parameters are simply the ratios of the FADS pressures, measured on the nose cap, to total pressure. Ideally, the parameters would not vary with flow angle. Thus, minimal and well-behaved variations of the parameter values with angle of attack were two criteria for selectivity (Larson et at., 1987). Figure 3 shows the variations of the three parameters with angle of attack for the five test Mach numbers. Local zero-flow angles were evidently obtained for PTPI4 and PTPIs, as indicated by the maxima for their curves in figures 3(a) and 30a), respectively. These maxima correspond within 1 percent of the theoretical ratios of stagnation to total pressure, which are shown in the legends of the figures. This agreement validates both the FADS pressure measurements and the wfmd-mnnelseference values of Mach number and total pressure.

For subsonic Mach numbers, a reasonably accurate value for stagnation pressure can be obtained without pre- cisely knowing Mach number or angle of attack. At low angles of attack, a value of 0.99 for PTP14 will suffice for computing stagnation pressure to within 1 percent of the actual value. For moderate angles of attack (100 to 20°), PTP15 would be the most accurate. As indicated by the curves' trends, and as verified by the full-scale flight tests (l..arson et at., 1987), PTP16 would be best at angles of attack higher than 20 °.

If corrections for Mach number and flow angles are applied, any of the three parameters could provide accurate determination of stagnation pressure for this study's test conditions. The requirement for Mach number and angle of attack implies that an iterative calculation scheme is necessary to determine airdata quantities from the FADS. A discussion of iteration methods is given in Larson and Siemers (1980).

The variation of static-pressure parameters for three locations with angle of attack is shown in figure 4. These parameters are the ratios of averaged FADS pressures to the free-stream static pressure. The static-pressure param- eters were determined from orifices on the fuselage nose because the pressures there are closer to static pressure than are those sensed on the nose cap. Ideally, the static-pressure parameters should not vary with Mach number or flow angle. However, as expected, the three static-pressure parameters are sensitive to Mach number and angle of attack. Because data were obtained only at five Mach numbers, accurate relationships between the parameters and Mach number cannot be established, as was accomplished in the flight tests which utilized data from accelerations and decelerations. But the effect of angle of attack on the static-pressure parameters were well defined because the data were obtained at 2 ° increments.

The variations of static-pressure parameters with angle of attack are dependent on Mach number, each parameter having a maximum variation of approximately 0.05 for particular test Mach numbers. Although the larger variations generally Occur at the higher Mach numbers, parameter PP32 also shows significant variations at the lower two test Mach numbers. The effect of Mach number on the magnitude of the static-pressure parameters can also be evaluated from figure 4. The values of the parameters increase with Mach number, with parameter PP22 having the largest change. The variation of the parameters with angle of sideslip (not shown) were a few percent also.

Figure 5 shows the variation of five angle-of-attack parameters with angle of attack. These parameters are the dif- ferences between the two FADS pressures divided by the free-stream dynamic pressure. Parameters AP11, API2, API3, and API4 use pressures sensed on the nose cap, while AP21 and AP31 use pressures sensed on the fuse- lage nose.

Parameter APl2 may be the best compromise of the angle-of-attack parameters. The reasons for this choice are (1) its large angle of attack sensitivity (slope of best straight line of parameter with angle of attack), (2) good linearity with angle of attack, and (3) minimum variation of the parameter with Mach number. The variation of the parameter's angle of attack sensitivity for the different Mach numbers ranges from 0.070/deg to 0.078/deg. These sensitivity values are much higher than corresponding sensitivities determined from the fuselage measurements, which vary from 0.022/deg to 0.043/deg. The curves for one of the fuselage parameters, AP21, show good linearity and almost no variation with Mach number. Thus, this parameter would also allow very accurate determination of angle of attack if precise measurements are made of the FADS pressures. The lower angle of attack sensitivity of parametcr AP21 dictates that more accuratc FADS pressure measurements must be obtained to achieve the same level of angle-of-attack accuracy as parameter API2.

It is concluded that with proper instrumentation and knowledge of Mach number, all the presented parameters could be used for accurate angle of attack determination. However, although not shown here, angle-of-attack errors as large as 1 ° could occur if sideslip effects are not accounted for when flying at angles of sideslip near 8°. This is in agreement with the results ofLarson et al. (1987).

Figure 6 shows the variation of four angle-of-sideslip parameters with angle of sideslip. These parameters are calculated in the same manner as the angle-of-attack parameters. Parameters BPll and BP12 use pressures sensed on the nose cap and the other parameters use pressures sensed on the fuselage nose. All these parameters, using the same orifice locations, were previously investigated in the flight tests reported in Larson et al. (1987), although that study's limited supersonic data were not published. Because the amount of wind-tunnel data taken at any particular angle of attack was small (see table 2), all of the sideslip parameter data (covering an angle-of-attack range from -4 ° to 18 °) are plotted in figure 6. Therefore, angle-of-attack effects, as well as Mach number effects, mask the linearity that would result if the data were plotted for a single angle of attack at constant Mach number.

The sensitivities of the sideslip parameters with angle of sideslip are highest for the nose cap orifices, as they were for the angle-of-attack parameters. The sensitivities for B Pll, for example, vary from approximately 0.062/deg to 0.072/deg, depending on the Mach number. These variations can be compared with those of the fuselage nose parameter, BP32, which has sensitivities that vary from 0.026/deg to 0.033/deg. The BP32 parameter is judged to be the most accurate of the fuselage nose parameters for ascertaining angle of sideslip.

Figure 7 shows the variation of three Mach number parameters with angle of attack. The first parameter (fig.

7(a)) is the ratio of two pressures, Pi and/94, both sensed on the nose cap. The other two parameters (figs. 7(b) and 7(c)) "also use ratios of two pressures. However, one pressure is sensed on the nose cap for a near-stagnation- pressure measurement, while the other pressure is sensed on the fuselage nose for a near-static-pressure measurement.

For Mach number determination from these parameters, large variations of their values with angle of attack are undesirable. However, all three parameters could be used for determining Mach number, if angle of attack is known, ralher than using separate measurements of corrected stagnation pressure and corrected static pressure.

Larson et al. (1987) presented nose cap Mach number parameters using pressure ratios of p3/P6 and p3/PT.

The angles of attack for those data were no greater than approximately 8°. The wind-tunnel data show that these parameters have high sensitivity to angle of attack above 8° . Thus, they are not shown here. The nose cap parameter with the lcast sensitivity to angle of attack is MPI. It utilizes pressure Pl which, because of an instrumentation problem, was not available in the flight program. However, even though MP1 is the least sensitive of the nose cap parameters to angle of attack, its sensitivity, especially at the higher Mach numbers, is larger than desirable. The other two Mach number parameters, which use fuselage pressures, have much less sensitivity with angle of attack and, as desired, have more variation with Mach number.

In selecting the most accurate parameter for Mach number determination, another factor should be considered in addition to desired sensitivity to Mach number and undesired sensitivity to flow angle. That factor is the importance of the relative magnitude of the pressures in the pressure ratio to the pressure transducer accuracy. However, since all the pressures used in the Mach number parameters were large compared to the accuracies of the pressure measure- mcnts, this factor was not significant for these tests. Therefore, parameter MP6 , which has the largest sensitivity to Math number and only moderate sensitivity to angle of attack, is the most accurate of the parameters studied for the wind-tunnel test conditions. However, there may be flight conditions for which the measured pressures could be low enough so that the magnitude of the pressures would be important for parameter selectivity.

CONCLUD G REMARKS

A flush airdata system on an F-14 model was tested in the 11- by ll-ft Unitary Wind Tunnel at the NASA Ames Research Center. Pressure orifices were located on a spherical surface of a test nose cap that was blended smoothly to the contours of the airplane's nose section. Orifices were also placed in two rows around the fuselage nose. Measurements were obtained at five Mach numbers from 0.73 to 1.39. Angles of attack were varied in 2 ° increments from -4* to 20 ° and angles of sideslip were varied in 4 ° increments from - 8 ° to 8".

Nose cap pressures that were investigated for stagnation-pressure determination agreed to within 1 percent of the tunnel-determined stagnation pressure when the local flow angles were zero. This helped validate the FADS pressure measurements as well as verify the wind-tunnel test conditions, especially at supersonic Mach numbers. It was concluded that any of the three pressures investigated could provide accurate determination of stagnation pressure.

The pressure parameters investigated for static-pressure determination were all sensitive to Mach number and angle ofattack. The variation of these parameters with angle of attack was well defined. However, the limited number of five test Mach numbers precluded accurately defining the variations with Mach number, as was accomplished in the flight tests that utilized accelerations and decelerations.

One angle-of-attack parameter was judged to be the best compromise between large angle-of-attack sensitivity, linearity with angle of attack, and small variability with Mach number. All of the parameters investigated, including those on the fuselage, could be used for accurate determination of angle of attack with proper instrumentation and knowledge of Mach number and dynamic pressure.

The angle-of-sideslip parameters performed in a manner similar to the angle-of-attack parameters. The sensitiv- ity of the parameters to changes in sideslip was highest for those using nose cap pressures.

Two types of Mach number parameters were investigated: one used the ratio of two nose cap pressures, while the other used the ratio of a nose cap pressure to simulate stagnation pressure and a fuselage pressure to simulate static pressure. The second type of parameter was better because although its sensitivity to angle of attack was less, at the same time its sensitivity to Mach number was much greater.

Ames Research Center Dryden Flight Research Facility National Aeronautics and Space Administration Edwards, California, March 31, 1989 REFERENCES Larson, Terry J., Stuart G. Flechner, and Paul M. Siemers Ill, Wind Tunnel Investigation of an All Flush Orifice Air Data System for a Large Subsonic Aircraft, NASA TP- 1642, 1980.

Larson, Terry J., and William G. Schweikhard, "Use of the Shuttle Entry Air Data Pressure System at Subsonic Speeds," Proc. Second Biennial Air Data Systems Conference, May 1978, pp. IV-8-1 to IV-8-7.

Larson, Terry J., and Paul M. Siemers III, "Subsonic Investigation of an All Flush Orifice Air Data System," Proc.

1980 A ir Data Systems Conference, Colorado Springs, Colorado, May 1980.

Larson, Terry J., and Paul M. Siemers III, Subsonic Tests of an All-Flush-Pressure-Orifice Air Data System, NASA TP-1871, 1981.

Larson, Terry J., and Paul M. Siemers III, Use of Nose Cap and Fuselage Pressure Orifices for Determination of Air Data for Space Shuttle Orbiter Below Supersonic Speeds, NASA TP-1643, 1980.

Larson, Terry J., Stephen A. Whitmore, L.J. Ehemberger, J. Blair Johnson, and Paul M. Siemers III, Qualitative Eval- uation of a Flush Air Data System at Transonic Speeds and High Angles of Attack, NASA TP-2716, 1987.

Research Facilities Handbook, NASA Ames Research Center, Moffett Field, Califomia, June 1982.

Siemers, P.M., III, and T.J. Larson, "Space Shuttle Orbiter and Aerodynamic Testing," J. Spacecraft and Rockets, vol. 16, no. 4, July-Aug. 1979, p. 223.

Siemers, P.M., III, H. Wolf, and M.W. Henry, "Shuttle Entry Air Data System (SEADS)--Flight Verification of an Advanced Air Data System Concept," AIAA-88-2104, May 1988.

Table 1. Wind-tunnel test data and computed FADS parameters,/3 = 0 °.

(a) Mach _ 0.73 Mach a p_ po _/ pl/pt p2/p_ p3/pt p4/pt ps/pt p6/pt pT/pt ps/pt pg/pt p_o/pt p_/pt p_2/pt p_s/p_ p_4/h p_s/pt p_6/pt p_7/_ p_s/pt p_9/_ p'zo/p_ p21/p_ p-_/pt p_/pt 0.735 -4.000 8.080 5.640 2.130 0.917 0.938 0.991 0.988 0.894 0.759 0.738 0.818 0.898 0.895 0.818 0.801 0.706 0.699 0.729 0.747 0.722 0.685 0.663 0.653 0.658 0.686 0.725 0.736 -2.000 8.080 5.639 2.140 0.904 0.924 0.984 0.996 0.916 0.787 0.760 0.829 0.903 0.903 0.828 0.789 0.706 0.710 0.737 0.736 0.719 0.692 0.672 0.660 0.667 0.690 0.721 0.736 0.100 8.078 5.634 2.140 0.888 0.906 0.975 0.999 0.934 0.814 0.780 0.835 0.907 0.909 0.835 0.777 0.705 0.719 0.742 0.728 0.716 0.696 0.682 0.666 0.677 0.695 0.718 0.735 2.000 8.077 5.642 2.130 0.872 0.889 0.963 0.998 0.948 0.838 0.799 0.842 0.910 0.909 0.839 0.766 0.704 0.729 0.744 0.718 0.712 0.700 0.689 0.673 0.685 0.699 0.714 0.733 4.000 8.075 5.649 2.130 0.855 0.870 0.948 0.998 0.964 0.861 0.819 0.845 0.909 0.9il 0.843 0.753 0.706 0.740 0.746 0.709 0.708 0.702 0.698 0.682 0.694 0.702 0.712 0.737 8.000 7.892 5.503 2.090 0.821 0.825 0.913 0.992 0.984 0.902 0.858 0.844 0.904 0.906 0.841 0.733 0.702 0.762 0.746 0.693 0.696 0.700 0.711 0.698 0.705 0.695 0.695 0.733 I0.000 7.896 5.522 2.080 0.803 0.802 0.89i 0.985 0.991 0.922 0.876 0.841 0.899 0.900 0.838 0.722 0.702 0.775 0.740 0.688 0.691 0.697 0.717 0.711 0.714 0.693 0.689 0.731 12.100 7.897 5.534 2.070 0.784 0.777 0.866 0.974 0.997 0.941 0.897 0.834 0.890 0.891 0.831 0.713 0.702 0.791 0.737 0.680 0.685 0.695 0.723 0.725 0.721 0.688 0.684 0.735 14.100 7.892 5.512 2.080 0.762 0.748 0.841 0.961 0.999 0.957 0.913 0.826 0.877 0.879 0.819 0.702 0.695 0.801 0.726 0.673 0.674 0.686 0.725 0.735 0.722 0.681 0.677 0.739 16.000 7.897 5.496 2.100 0.742 0.719 0.812 0.945 0.996 0.967 0.927 0.813 0.865 0.865 0.805 0.689 0.688 0.813 0.716 0.667 0.661 0.674 0.726 0.742 0.724 0.671 0.667 0.736 19.900 7.900 5.513 2.090 0.708 0.670 0.755 0.910 0.990 0.986 0.952 0.787 0.837 0.834 0.773 0.674 0.684 0.838 0.694 0.662 0.649 0.656 0.732 0.768 0.732 0.655 0.654 Table 1. Continued.

(b) Mach _ 0.90 Mach _ Pt P, _ pl/pt P2/Pt P3/pt p4/Pt Ps/Pt p6/Pt p7/pt Ps/P_ Pg/Pt Pl0/Pt Pll/pt PI2/Pt PI3/Pt pI4/Pt plS/Pt pl6/pt PI?/Pt plS/Pt PI9/Pt P20/Pt P21/Pt P'22 ]Pt P23 ]Pt 0.899 -4.000 7.404 4.383 2.480 0.888 0.916 0.985 0.988 0.874 0.700 0.667 0.768 0.872 0.871 0.770 0.735 0.605 0.606 0.644 0.659 0.629 0.580 0.549 0.533 0.541 0.582 0.630 0.900 -2.000 7.404 4.377 2.480 0.873 0.900 0.977 0.994 0.897 0.735 0.693 0.781 0,878 0.878 0.781 0.721 0.606 0.619 0.654 0.645 0.624 0.589 0.563 0.542 0.555 0.586 0._24 0.900 0.000 7.403 4.377 2.480 0.852 0.878 0.966 0.998 0.917 0.763 0.717 0.789 0.882 0.881 0.786 0.704 0.605 0.632 0.659 0.632 0.617 0.592 0.572 0.551 0.565 0.593 0.620 0.900 2.000 7.410 4.381 2.480 0.832 0.855 0.950 0.999 0.937 0.793 0.740 0.795 0.883 0.886 0.794 0.690 0.605 0.645 0.664 0.618 0.612 0.596 0.583 0.560 0.575 0.595 0.615 0.900 4.000 7.434 4.395 2.490 0.814 0.831 0.934 0.998 0.954 0.822 0.767 0.798 0.883 0.884 0.796 0.674 0.606 0.656 0.667 0.609 0.607 0.599 0.593 0.570 0.586 0.597 0.610 0.903 6.000 7.399 4.362 2.490 0.792 0.803 0.913 0.995 0.969 0.849 0.792 0.800 0.881 0.885 0.796 0.660 0.606 0.673 0.667 0.597 0.600 0.599 0.601 0.582 0.595 0.593 0.600 0.902 8.000 7.400 4.368 2.490 0,769 0.776 0.891 0.990 0.980 0.875 0.814 0.797 0.878 0.880 0.793 0.645 0.604 0.687 0.663 0.584 0.590 0.595 0.611 0.595 0.605 0.593 0.597 0.901 10.000 7.401 4.370 2.480 0.744 0.745 0.863 0.981 0.989 0.899 0.840 0.793 0.870 0.872 0.788 0.632 0.604 0.703 0.660 0.576 0.583 0.594 0.620 0.613 0.612 0.590 0.586 0.896 12.000 7.403 4.397 2.470 0,724 0.716 0.838 0.972 0.996 0.922 0.863 0.790 0.861 0.864 0.784 0.622 0.602 0.723 0.655 0.573 0.577 0.591 0.628 0.627 0.623 0.584 0.579 0.901 14.000 7.404 4.372 2.490 0.701 0.686 0.808 0.957 0.998 0.941 0.883 0.779 0.849 0.852 0.774 0.608 0.594 0.737 0.646 0.564 0.565 0.580 0.633 0.641 0.627 0.576 0.568 0.902 16.100 7.406 4.368 2.490 0.678 0.652 0.775 0.941 0.998 0.956 0.902 0.767 0.836 0.838 0.759 0.597 0.588 0.756 0.635 0.557 0.553 0.571 0.639 0.658 0.635 0.565 0.560 0.901 18.000 7.404 4.373 2.480 0.658 0.622 0.744 0.922 0.997 0.970 0.920 0.753 0.822 0.821 0.745 0.589 0.590 0.771 0.622 0.555 0.547 0.563 0.646 0.676 0.640 0.555 0.554 0.901 20.000 7.397 4.368 2.480 0.639 0.591 0.709 0.902 0.993 0.981 0.937 0.738 0.804 0.804 0.727 0.578 0.574 0.788 0.610 0.551 0.537 0.550 0.650 0.693 0.645 0.543 0.547 Table 1. Continued.

(c) Mach _ 1.05 Mach cx pt 10# q 101/11_ p2/Pl p3/Pt lO4/lOt 1_/lOt P6/19t i07/ll)t p$/10t Pg/Pt PlO/Pt pll ]Pl PI2]Pl p13 ]il_ p14 ]PI PI5/IDI p16 ]iI_ pIT/iDt pl$/Pt P19/Pt p'20 ]Pt p21 [Pt p2.2/Pt P"23 ]Pt 1.048 0.000 7.403 3.696 2.840 0.833 0.863 0.961 0.997 0.912 0.741 0.687 0.766 0.871 0.871 0.765 0.667 0.518 0.587 0.618 0.575 0.560 0.535 0.515 0.487 0.504 0.532 0,560 1.044 2.000 7.401 3.711 2.830 0.813 0.836 0.944 0.999 0.932 0.774 0.712 0.771 0.871 0.873 0.768 0.650 0.520 0.600 0.622 0.561 0.553 0.539 0.523 0.497 0.515 0.535 0.557 1.055 4.000 7.411 3.669 2.860 0.787 0.807 0.925 0.996 0.948 0.801 0.735 0.77! 0.8_8 0.870 0.769 0.630 0,515 0.610 0.619 0.539 0.540 0.534 0.527 0.501 0.519 0.532 0.545 1.050 6.000 7.404 3.688 2.840 0.768 0.782 0.905 0.996 0.963 0.831 0.764 0.776 0.869 0.871 0.771 0.617 0.520 0.630 0.623 0.533 0.536 0.537 0.540 0.518 0.533 0.529 0.539 1.053 8.000 7.403 3.671 2.850 0.738 0.748 0.876 0.989 0.976 0.860 0.790 0.771 0.861 0.864 0.769 0.599 0.516 0.645 0.617 0.513 0,522 0.532 0.548 0.531 0.541 0.523 0.525 1.042 10.100 7.403 3.721 2.830 0.717 0.720 0.853 0.980 0.987 0.888 0.819 0.770 0.856 0.859 0.766 0.585 0.519 0.667 0.617 0.508 0.517 0.532 0.559 0.551 0.553 0.525 0.522 1.034 12.050 7.401 3.755 2.810 0.695 0.690 0.825 0.969 0.993 0.910 0.843 0.766 0.849 0.850 0.761 0.577 0.519 0.687 0.613 0.502 0.508 0.530 0.572 0.570 0.565 0.520 0.514 1.051 14.100 7.408 3.683 2.850 0.667 0.651 0.792 0.956 0.998 0.928 0.864 0.754 0.835 0.837 0.747 0.555 0.508 0.698 0.599 0.482 0.489 0.511 0.572 0.581 0.565 0.497 0.490 1.052 16.000 7.418 3.683 2.850 0.645 0.619 0.762 0.941 0.999 0.947 0.883 0.742 0.824 0.824 0.735 0.543 0.497 0.716 0.589 0.473 0.476 0.500 0.576 0.596 0.571 0.483 0.473 1.046 18.100 7.439 3.721 2.850 0.623 0.585 0.729 0.920 0.998 0.963 0.904 0.730 0.808 0.809 0.720 0.535 0.487 0.737 0.578 0.467 0.464 0.491 0.587 0.617 0.581 0.475 0.460 1.029 20.000 7.407 3.780 2.800 0.608 0.556 0.694 0.899 0.995 0.976 0.924 0.719 0.794 0.792 0.704 0.530 0.497 0.758 0.569 0.466 0.457 0.483 0.598 0.644 0.595 0.468 0.455 Table1. Continued.

(d) Mach_ 1.20

Mach c, Pt P., _ pl/pt p2/Pt p3/Pt p4/p_ ps/pt p6/iot pT/pt ps/pt Pg/Pt I)lo/Pt pit/pt pl 2/'Pt pl3/Pt ]014/pt PlS/Pt pl6/Pt pl7[Pt pls/Pt pl9/Pt p20 [Pt P21 [Pt p22 [Pt p23/Pt 1.201 -4.000 7.401 3.050 3.080 0.828 0.867 0.966 0.980 0.845 0.618 0.549 0.687 0.826 (I.826 0.684 0.638 0.435 0.487 0.537 0.534 0.507 0.451 0.418 0.396 0.407 0.451 0.502 1.201 -2.000 7.399 3.048 3.080 0.803 0.840 0.953 0.987 0.870 0.652 0.574 0.694 0.831 0.830 0.691 0.617 0.440 0.499 0.543 0.517 0.498 0.456 0.428 0.404 0.419 0.455 0.496 1.203 0.000 7.400 3.042 3.080 0.778 0.813 0.939 0.990 0.891 0.685 0.603 0.701 0.833 0.832 0.698 0.601 0.443 0.513 0.550 0.501 0.490 0.461 0.439 0.414 0.433 0.458 0.487 1.202 2.000 7.400 3.045 3.080 0.753 0.784 0.919 0.991 0.913 0.717 0.631 0.705 0.834 0.835 0.701 0.583 0.446 0.529 0.554 0.484 0.482 0.465 0.452 0.426 0.445 0.459 0.478 I. 198 4.000 7.401 3.059 3.070 0.726 0.751 0.898 0.988 0.931 0.751 0.663 0.708 0.832 0.834 0.705 0.565 0.447 0.545 0.555 0.469 0.474 0.465 0.462 0.438 0.458 0.462 0.473 1.195 6.000 7.401 3.070 3.070 0.701 0.721 0.874 0.986 0.950 0.786 0.698 0.714 0.830 0.834 0.708 0.548 0.448 0.565 0.556 0.459 0.466 0.466 0.475 0.454 0.470 0.462 0.467 1.190 8.000 7.402 3.094 3.060 0.675 0.688 0.848 0.980 0.964 0.821 0.734 0.715 0.828 0.829 0.707 0.532 0.449 0.586 0.558 0.446 0.459 0.467 0.487 0.473 0.485 0.460 0.457 1.197 10.000 7.437 3.078 3.090 0.643 0.647 0.817 0.971 0.973 0.844 0.758 0.705 0.815 0.820 0.700 0.510 0.432 0.600 0.546 0.429 0.441 0.459 0.494 0.486 0.491 0.451 0.440 1.194 12.000 7.410 3.079 3.070 0.618 0.613 0.789 0.959 0.982 0.873 0.788 0.701 0.810 0.814 0.697 0.501 0.435 0.624 0.545 0.422 0.432 0.455 0.505 0.505 0.502 0.444 0.431 1.186 14.100 7.396 3.106 3.060 0.598 0.582 0.759 0.946 0.989 0.903 0.822 0.702 0.802 0.806 0.693 0.491 0.437 0.646 0.540 0.414 0.425 0.449 0.517 0.525 0.5i2 0.440 0.425 1.176 16.000 7.400 3.149 3.050 0.582 0.553 0.732 0.931 0.991 0.926 0.851 0.698 0.795 0.797 0.689 0.483 0.439 0.669 0.532 0.411 0.415 0.444 0.528 0.546 0.525 0.431 0.415 1.163 18.050 7.400 3.202 3.030 0.566 0.521 0.700 0.914 0.993 0.945 0.876 0.688 0.782 0.784 0.680 0.476 0.441 0.693 0.524 0.407 0.408 0.437 0.540 0.571 0.535 0.424 0.412 1.144 20.100 7.401 3.278 3.010 0.559 0.499 0.674 0.894 0.991 0.%5 0.905 0.684 0.774 0.775 0.673 0.477 0.448 0.723 0.522 0.410 0.404 0.433 0.555 0.599 0.551 0.420 0.404 Table1. Concluded.

(e) Mach _ 1.39 Mach ot Pt lo. q p,/Pt p2/pt P3/Pt io4/pt Ps/Pt p6/Pt pT/_ Ps/Pt Pg/Pt plo/r_ p, lr_ pnlr_ pl31r_ p_4/_ p_slr_ p161pt p_'71r_ p_8/pt p_91_ l:,2o/pt rntlpt r,'_2/_ _3/pt 1.389 -4.000 7.388 2.358 3.180 0.718 0.768 0.912 0.950 0.807 0.552 0.368 0.578 0.760 0.754 0.561 0.548 0.335 0.392 0.447 0.451 0.424 0.364 0.334 0.313 0.325 0.364 0.413 1.390 -2.000 7.391 2.357 3.190 0.689 0.739 0.898 0.955 0.830 0.577 0.394 0.575 0.759 0.758 0.564 0.528 0.338 0.404 0.452 0.434 0.418 0.372 0.345 0.321 0.337 0.368 0.407 1.389 0.000 7.388 2.358 3.180 0.656 0.703 0.877 0.956 0.850 0.601 0.427 0.574 0.757 0.758 0.563 0.504 0.339 0.418 0.454 0.417 0.406 0.375 0.358 0.332 0.350 0.374 0.401 1.388 2.000 7.389 2.361 3.180 0.623 0.667 0.855 0.953 0.867 0.629 0.467 0.573 0.753 0.756 0.565 0.486 0.342 0.432 0.458 0.399 0.395 0.376 0.367 0.343 0.361 0.375 0.393 1.387 4.000 7.389 2.364 3.180 0.593 0.636 0.836 0.956 0.886 0.658 0.507 0.573 0.752 0.756 0.568 0.471 0.349 0.449 0.460 0.385 0.388 0.379 0.378 0.355 0.372 0.374 0.383 1.387 6.000 7.393 2.365 3.190 0.555 0.594 0.808 0.951 0.903 0.693 0.551 0.573 0.747 0.753 0.569 0.451 0.349 0.471 0.463 0.371 0.377 0.379 0.390 0.369 0.385 0.371 0.375 1.385 8.000 7.395 2.374 3.190 0.519 0.556 0,780 0.944 0.919 0.725 0.593 0.574 0.743 0.748 0.568 0.431 0.351 0.487 0.460 0.355 0.365 0.376 0.400 0.386 0.398 0.371 0.366 1.381 10.150 7.402 2.388 3.190 0.484 0.517 0.753 0.938 0.934 0.758 0.632 0.573 0.738 0.744 0.569 0.416 0.349 0.512 0.457 0.346 0.355 0.372 0.413 0.404 0.411 0.364 0.353 1.380 12.050 7.403 2.393 3.190 0.449 0.487 0.728 0.926 0.942 0.787 0.667 0.573 0.734 0.739 0.569 0.402 0.350 0.531 0.451 0.335 0.345 0.366 0.423 0.421 0.420 0.357 0.344 1.375 14.150 7.402 2.408 3.190 0.415 0.453 0.699 0.913 0.948 0.817 0.704 0.572 0.725 0.732 0,567 0.387 0.351 0.554 0.442 0.324 0.332 0.359 0.430 0.439 0.427 0.354 0.334 1.371 16.100 7.398 2.421 3.190 0.398 0.425 0.673 0,901 0.956 0.846 0.739 0.573 0.721 0.723 0.568 0.379 0.356 0.577 0.437 0.318 0.322 0.354 0.440 0.459 0.439 0.346 0.322 1.365 18.100 7.402 2.444 3.190 0.384 0.401 0.646 0.886 0.959 0.872 0.772 0.572 0.713 0.716 0.565 0.368 0.364 0.601 0.428 0.311 0.308 0.343 0.451 0.480 0.450 0.339 0.307 1.357 20.000 7.401 2.470 3.180 0.376 0.380 0.617 0.867 0.960 0.895 0.805 0.572 0.706 0.707 0.564 0.361 0.370 0.626 0.421 0.309 0.298 0.337 0.460 0.502 0.463 0.335 0.299 g.

o _ _ r _.

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I I I I I I I ORIGINAE PAGE ORIGINAL PAGE I_ BLACK AND WHITE PHOTOGRAPH OF POOR QUALITY 905B Figure 1. F-14 wind-tunnel model.

1, ORIGINAL pAGE IS OF pOOP. QUALIW Orifices 1 to 7 are spaced in increments o1 20 ° Orifices 8 to 11 are spaced In increments of 30 ° All orifices are 0.034 cm in

, .S98-cmzdi s diameter

• \ |ongitudinal _ 5_ xis _- Model / Front view Side view All orifices are 0.034 cm in diameter Rows are normal to fuselage longitudinal axis Station 61 orifices are on vertical and horizontal centerlines Station 127 orifices are spaced 45 ° apart 12 16 15 13 22 18 14 20 Station 61 Station 127 Figure 2. Locations of FADS nose cap and fuselage stalion orifices.

Pl2/Pl Mach at _ = 2 ° 1.05 I 0 0.73 [] 0.90 1.00 1.05 0.999 1.20 0.993 1.39 0.961 PTP14 .95 Nose cap .90

85 ! I I

I

-5 0 5 10 15 2O (I, deg 996O (a) Parameters PTPI4.

1.05 PI2/Pl Mach at (7. : 18"

F

0 0.73 -- I-I 0.90 1 1.1111 1.05 0.999 1.20 0.993 I_ 1.39 0.961 .95 PTP15 .90 .85 = lw

I

80 i 1 I

-5 0 5 10 15 20 _., de9 9cJI61 (b) Parameters PTPI5.

Figure 3. Variation of stagnation pressure parameters with angle of altack.

2O 1.0 Mach O 0.73 rl o.9o .9 1.05 1.20 D 1.39 PTP16 Nose cap 0 5 10 15 20 25 C/, deg (c) Parameters PTP16.

Figure 3. Concluded.

1.5 Mach 0 0.73 [3 0.90 1.4

0 .os

Z_ 1.20 1.39 1.3 PP22 1.2 Station 61 1.1 1.0 -5 0 5 10 15 20 25 (_, deg (a) Parameter PP22.

Figure 4. Variation of static pressure parameters with angle of attack.

2!

Mzch 0 0.7_ CZ O.90 I.0S I_3 1.20 D 1.39 Station 61 PP32 1.1 1.0 _ "" v .9 _ 5 10 -5 _, deg (b) parameter PP3_.

Mach 0 o.73 o.go _, 1.05 1.3 _, 1.20 1.39 P23_ 17 1.2 P2 _ P19 Station 137

P%3 1.1 - ",," O' __

1.0 0 25 .9 5 ms -5 0 (X, deg (c) Parameter PP33.

Figure 4. Concluded.

.8 Mach O 0.73 [] 0.90

<> 1.os

/_ 1.2o .4 17 1.39 Nose cap AP11 -.4 -.8 -1.2

I I 1 I

-1.6 -5 5 10 15 20 25 (x, deg 9966 (a) Parameter APll.

Figure 5. Variation of angle of attack parameters with angle of attack.

1.2 Mach O o.73 D O.9O I.05 1.20 .8 D 1.39 .4 Nose cap AP12 -.4 -.8

I

I l I I !

-1.2 15 20 -5 0 5 10 (I, deg _7 (b) Parameter AP12.

Figure 5. Continued.

1.2 Mach O 0.73

r

[] 0.90 1.05 Z_ 1.20 .8 [7 1.39 .4 Nose cap AP13 -.4 -.8

I I

-1.2 -5 20 25 (c) Parameter AP13.

Figure 5. Continued.

Math O 0.73 E] 0.90 1.05 Z_ 1.20 P 1.39 A P14 Nolle cap

I I I i

I I

0 5 10 15 2O 0_, deg (d) Parameter API4.

m .8 Mach 0 0.73 El 0.90 1.05 1.20 ,4 [7 1.39 AP21 0 Station 01 -.4 1 1 I 1 I I -.8 -5 0 5 10 15 20 25 (1, deg 9970 (c) Parameter AP21.

Figure 5. Continued.

Mach 0 0.73 [] 0.90 1.05 Z_ 1.20 .4 E_ 1.39 AP31 Station 127 I I I I I -.4 -5 10 15 20 25 0 5 CX, deg (0 Parameter AP3].

Figure 5. Concluded.

.8 Mach O 0.73 [] 0.90 1.05 .6 Z_ 1.20 I_ 1.39 Pl 1 .4

|

Nose cap BP11

J

0 m -°2 B

i

°.4 B I I ! I I _ I I I I -.6 -10 -8 -6 -4 -2 0 2 4 6 8 13, deg (a) Parameter BPll.

Figure 6. Variation of angle of sideslip parameters with angle of sideslip.

.8 -- .6 C 0.73 D 0.90 1.05 1.20 P 1.39 .4 Pl0

|

Nose cap BP12 0 -.2

e

-.4 I ! 1 I I -.6 ! i I I I -10 -8 -6 -4 -2 0 2 4 6 8 10 (b) Paramctcr BPnz.

Mach 0 0.73 rn 0.90 O 1.05 L_ 1,20 D 1.39 .6 -- .4 BP21 .2 o qw

! I I ! I ! i I t]

I -.2 -10 -8 -6 -4 -2 0 2 4 6 8 9, deg (c) Parameter BP2x.

Figure 6. Continued.

Mach O 0.73 [] 0.90 P23_P17 <> 1.05 /% 1.20 P21 _ P19 I:P 1.39 Station 127 .4 4t, BP32 0 O -.2 I I I I 1 I I -.4 I I I - 10 -8 -6 -4 -2 0 2 4 6 8 10 13, deg (d) Parameter BP32.

Figure 6. Concluded.

1.0 Mach O 0.73 [] 0.90 1.05 1.20 D 1.39 MP1 .8 .6 Nose cap I .4 -5 0 5 10 15 20 _, deg (a) Parameter MPI.

Figure 7. Variation of Mach number parameters with angle of attack.

Mach 0 0.73 0.90 1.05 1.20 1.39 2.4 _ 2.2 No_ cap 2.0 MP 5 Station 61 1.8 1.6 O 1.4 C_C Occ I J 20 25 1.2 -_ 10 15 -5 0 5 (I, deg (b) Parameter MPs.

Figure 7. Continued.

3O Mach 0 o._3 t_ 0.9o 2.8 1.05 1.20 I_ 1.39 2,6 P4_ 2.4 Nose cat> 2.2 Station 127 MP 6 2.0 1.8 E_ _ _ ,_ _ .g--_ _ __-_-_ 1 °6

,..,CO

1,4 1.2 _ 10 15 0 5 997s -5 G, deg (c) Parameter MF6, Figure 7. Concluded.

i1=1

Report Documentation Page

2. Government Accession No. 3. Recipient's Catalog No.

1. Report No.

NASA TM- 101697 5. Report Date 4. Title and Subtitle April 1990 Wind-Tunnel Investigation of a Rush Airdata System at Mach Numbers 6. Performing Organization Code From 0.7 to 1.4 8. Performing Organization Report No.

7. Author(s) H-1544 Terry L Larson, T'Lrnothy R. Mocs, and Paul M. Sicmers III 10. Work Unit No.

RTOP 505-6021 9. Performing Organization Name and Address 11. Contract or Grant No.

NASA Ames Research Center Dryden Right Research Facility EO. Box 273, Edwards, CA 93523-0273 13. Type Of r Report and Period Covered 12. Sponsoring Agency Name and Addreu Technical Memorandum 14. Sponsoring Agency Code National Aeronautics and Space Administration Washington, DC 20546 15. Supplementary Notes Terry J. Larson is affiliated with Analytical Mechanics Associates; Paul M. Siemers I]I is affiliated with Langley Research Center.

16. Abstract Flush pressure oririces installed on the nose section of a I/7-scale model of the F- 14 airplane were evaluated for use as a flush airdata system (FADS). Wind-tunnel tests were conducted in the 11- by 11-ft Unitary W'md Tunnel at NASA Ames Research Center. A fltll-scale FADS of the same configuration was previously tested using an F- 14 aircraft at the Dryden Right Research Facility of NASA Ames Research Center (Ames- Dryden). These tests, which have been published, are part of a NASA program to assess accuracies of FADS for use on aircraft. The test program also provides data to validate algorithms for the shuttle entry airdata system developed at the NASA Langley Research Center. The wind'tunnel test Mach numbers were 0.73, 0.90, 1.05, 1.20, and 1.39. Angles of attack were varied in 2* increments from -4 ° to 20 °. Sideslip angles were varied in 4 ° increments from -8 ° to 8°. Airdata parameters were evaluated for determination of free- stream values of stagnation pressure, static pressure, angle of attack, angle of sideslip, and Mach number.

These parameters are, in most cases, the same as the parameters investigated in the flight test program. The basic FADS wind-tunnel data are presented in tabular form. A discussion of the more accurate parameters is included.

18. Distribution Statement 17. Key Words (Suggested by Author(,=)) Unclassified -- Unlimited Airdata Flow-angle measurements Rush airdata system Subject category 06 Rush pressure orifices 21. No. of pages 22. Price 20. Security Classif. (of this page) 19. Security Classif. (of this report) 35 A03 Unclassified Unclassified NASA FORM1826 OCT 86 *For sale by the National Technical Information Service, Springfield, VA 22161-2171.