1、IASA TECHNICALREPORTNASA TR R-474REDICTION OF STATIC AERODYNAMICHARACTERISTICS FOR SLENDER BODIES.LONE AND WITH LIFTING SURFACESO VERY HIGH ANGLES OF ATTACKeland Howard Jorgensen.mes Research Centertoffett Field, Calif. 94035ITIONAL AERONAUTICSAND SPACE ADMINISTRATION WASHINGTON, D. C. SEPTEMBER1977
2、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. Report No. 2. Government Accession No. 3. Recipients Catalog No.NASA TR R-4744. Title and SubtitlePRE
3、DICTION OF STATIC AERODYNAMIC CHARACTERISTICSFOR SLENDER BODIES ALONE AND WITH LIFTING SURFACESTO VERY HIGH ANGLES OF ATTACK*7. Author(s)Leland Howard Jorgensen9. Performing Organization Name and AddressAmes Research CenterMoffett Field, California 9403512. Sponsoring Agency Name and AddressNational
4、 Aeronautics and Space AdministrationWashington, D. C. 205465. Report DateSeptember 19776. Performing Organization Code8. Performing Organization Report No.A-696810. Work Unit No.505-06-9711, Contract or Grant No,13. Type of Report and Period CoveredTechnical Report14, Sponsoring Agency Code15. Supp
5、lementary Notes“Formerly issued as NASA TM X-73,123 with limited distribution. Reissued for unlimited distribution.16. AbstractAn engineering-type method is presented for computing normal-force and pitching-moment coefficients forslender bodies of circular and noncircular cross section alone and wit
6、h lifting surfaces. In this method, asemi-empirical term representing viscous-separation crossflow is added to a term representing potential-theorycrossflow.For nmqy bodies of revolution, computed aerodynamic characteristics are shown to agree with measuredresults for investigated free-stream Mach n
7、umbers from 0.6 to 2.9. The angles of attack extend from 0 to 180 for Moo = 2.9 and from 0 to 60 for Moo = 0.6 to 2.0.For several bodies of elliptic cross section, measured results are also predicted reasonably well over theinvestigated Mach number range from 0.6 to 2.0 and at angles of attack from
8、0 to 60 . As for the bodies ofrevolution, the predictions are best for supersonic Mach numbers.For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-forcecoefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. Howeve
9、r, with adecrease in Mach number to Moo = 0.6, the agreement for CN rapidly deteriorates, although the normal-forcecenters remain in close agreement.Vapor-screen and oil-flow pictures are shown for many body, body-wing, and body-wing-tail configurations.When separation and vortex patterns are asymme
10、tric, undesirable side forces are measured for the models evenat zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase inMach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body crosssection (particula
11、rly the body nose).-i7. Key Words (Suggested by Author(s)High angle of attackNoncircular bodiesBodies of revolutionWing-body configurationsWing-body-tail configurations19, Security Classif. (of this reportUnclassifiedAerodynamic theoryVapor-screentechniqueVortex flowOil-flow technique18, Distributio
12、n StatementUnlimited20. Security Cla_if. (of this page)UnclassifiedSTAR Category - 0221, No. of Pages243“For sale by the National Technical Information Service, Springfield, Virginia 2216122. Price*$7.50Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,
13、-To the memory of H. Julian Allen,a giant in aerodynamic research.Born: 1 April 1910Died. 29 January 1977Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TABLE OF CONTENTSNOTATION .SUMMARY .CHAPTER 1. INTRODUCTION CHAPTER 2. DERIVATION OF BASIC METHOD
14、 FOR COMPUTINGCN AND Cm CHARACTERISTICS .2.1 Bodies of Revolution .2.2 Bodies of Circular and Noncircular Cross Section Alone andWith Lifting Surfaces .2.3 Empirical Input Values 2.3.1 Crossflow drag coefficient 2.3.2 Crossflow drag proportionality factor .2.4 Formulas and Values of (Cn/Cno)sB and (
15、Cn/Cno)_ewt_, for VariousCross Sections 2.4.1 Formulas of (Cn/Cno)sB 2.4.2 Formulas of (Cn/Cno)Newt 2.4.3 Values of (Cn/Cno)sB and (Cn/Cno)Newt 2.5 Relative Influence of Crossflow Terms CHAPTER 3. METHOD APPLIED TO BODIES OF REVOLUTION 3.1 Cone-Cylinder and Ogive-Cylinder Bodies at Moo = 2.9 3.20giv
16、e-Cylinder Bodies at Moo = 0.6 to 2.0 3.3 Predicted Effect of Change in Crossflow Reynolds Number fromSubcritical to Supercritical CHAPTER 4. METHOD APPLIED TO BODIES OF ELLIPTICCROSS SECTION 4.1 Bodies Studied and Tests at Moo = 0.6 to 2.0 4.2 Equations Used to Compute CN and Cm for Each Body .4.2.
17、1 Equations for bodies with constant a/b cross sections(bodies Bi and B2) .4.2.2 Equations for body with variable a/b cross sections(body B3 ) 111PageviI266ll1414171818192021232324262828293131Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4.3 Compar
18、isonof ComputedWith MeasuredNormal-ForceandNormal-Force-CenterCharacteristics.4.3.i Bodieswith constanta/b cross sections (bodies B1 and B2 ) .4.3.2 Body with variable a/b cross sections (body B3) .CHAPTER 5. METHOD APPLIED TO BODY-WING AND BODY-WING-TAILCONFIGURATIONS .5.1 Configurations Studied an
19、d Tests at Moo = 0.6 to 2.0 .5.2 Methodology Used to Compute CN and Cm 5.3 Comparison of Computed with Measured Normal-Force andNormal-Force-Center Characteristics .5.3.1 Body-wing configurations 5.3.2 Body-wing-tail configurations .CHAPTER 6. VISUAL OBSERVATION OF FLOWS OVER MODELS .6.1 Models Cons
20、idered 6.2 Vapor-Screen and Oil-Flow Techniques 6.2.1 Vapor-screen technique .6.2.2 Oil-flow technique .6.3 Photographs Obtained from Vapor-Screen and Oil-Flow Techniques.6.3.1 Photographs from vapor-screen technique 6.3.2 Photographs from oil-flow technique CHAPTER 7. EXPERIMENTAL SIDE FORCES ON MO
21、DELS AT 13 = 07.1 Bodies Alone 7.1.1 Effects of nose fineness ratio and Mach number .7.1.2 Effect of nose-tip rounding .7.1.3 Effect of afterbody side strakes 7.1.4 Effect of elliptic cross section 7.2 Body-Wing and Body-Wing-Tail Configurations .7.2.1 Effects of adding a wing and a wing plus tail t
22、o a body 7.2.2 Effects of wing aspect ratio and taper ratio .CHAPTER 8. CONCLUDING REMARKS Page333333353536393941434445454647485252545454545555555556ivProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-APPENDIX A DERIVATION OF (Cn/Cno)_ewt_ FOR WINGED-C
23、IRCULARAND WINGED-ELLIPTIC CROSS SECTIONS .Winged-Circular Cross Section with Wing Planform Perpendicular toCrossflow Velocity Winged-Elliptic Cross Section with Semimajor Axis and Wing PlanformPerpendicular to Crossflow Velocity .Winged-Elliptic Cross Section with Semiminor Axis and Wing PlanformPe
24、rpendicular to Crossflow Velocity .APPENDIX B - DERIVATION OF (Cn/Cno)_pwt.,_ FOR WINGED-SQUARECROSS SECTIONS WITH ROUNDED CORNERS .APPENDIX C - FORMULAS TO COMPUTE GEOMETRIC PARAMETERSFOR TANGENT OGIVES REFERENCES .FIGURES .Page59596O6264676875Provided by IHSNot for ResaleNo reproduction or network
25、ing permitted without license from IHS-,-,-NOTATION/1.Ih“II):Ir/1 Haa. t)(A( _1H)(/.(“m(.V( 1/(),(yd/“,/p/rbody cross-sectional areabody basearcalat x- _)planform areareference area _taken as A b for the comparisons of computed with experimentalresults)surface wetted areaexposed wing planform area (
26、2 panels)specd of soundscmimajor and semiminor axes of elliptic cross sectionF.axial-force coefficient,qooAr F.crossflow drag coefficient of circular cylinder section, - .qnA_ v )dc_d ra gdrag coefficient,qoo,t rLiftlift coefficient,qooArI)itching-lnoment coefficient about station at x m6,llormal-fl
27、)rce coefficient, - -q_oA rlocal normal-force coefficient per unit lengthP - Poepressure coefficient. -.-qoosidc-force coefficient, Fvq ooA rbody cross-section diameterpitching momentfromnose, - qooA rXcross force per unit length along body lengthpotential cross force per unit length along body leng
28、thviscous cross force per unit length along body lengthviProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-kKI ,K2_29_AM nMooPP_qnq_rraSstReRe nVv_Waxial, normal, and side forcercorner rounding for body cross section, -wlongitudinal and transverse appa
29、rent mass coefficientsbody lengthbody aftersection lengthbody nose lengthMach number component normal to body axis, Moo sin afree-stream Mach numberpressurefree-stream static pressuredynamic pressure component normal to body axis, qoo sin 2 xIfree-stream dynamic pressure, _ oV_oobody cross-section r
30、adius or corner radiusarc radius of ogivedisplacement of crossflowsemispantimeovooxfree-stream Reynolds number,- /adReynolds number component normal to body axis, Re _ sin abody volumevelocity component normal to body axis, Voo sin atree-stream velocitybody widthviiProvided by IHSNot for ResaleNo re
31、production or networking permitted without license from IHS-,-,-XxXacXcxDlXsfPSubscriptsbNewtHose0SBstagreference length (generally taken as d for the comparisons of computed withexperimental results)axial distance from body noseaxial distance from body nose to aerodynamic normal-force center (cente
32、r ofpressure)axial distance from body nose to centroid of body planform areaaxial distance frona body nose to pitching-moment reference centeraxial distance from body nose to aerodynamic side-force centerangle of attackangle of sideslipwing planform semiapex anglecrossflow drag proportionality facto
33、r (ratio of crossflow drag for a finite-lengthcylinder to that for an infinite-length cylinderjviscosity coefficient of airdensity of airangle of roll about body longitudinal axisbody basecylinderNewtonian theorybody noseequivalent circular body or cross sectionslender-body theorystagnationVUlProvid
34、ed by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-PREDICTION OF STATIC AERODYNAMIC CHARACTERISTICSFOR SLENDER BODIES ALONE AND WITH LIFTINGSURFACES TO VERY HIGH ANGLES OF ATTACK*Leland Howard JorgensenAmes Research CenterSUMMARYAn engineering-type method is
35、presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone andwith lifting surfaces. In this method, a semiempirical term representing viscous-separationcrossflow is added to a term representing potential-theory crossflow.
36、 111computing Qy andC m for bodies alone, slender-body theory is used for the term representing the potentialcrossflow. For bodies with thin wings and tails, the linearized potential method of Nielsen,Kaattari, and Pitts, modified for high angles of attack, is used.For many bodies of revolution, com
37、puted aerodynamic characteristics are shown toagree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. Theangles of attack extend froin 0 to 180 for Moo = 2.9 and from 0 to 60 for Moo = 0.6to 2.0.For several bodies of elliptic cross section, measured results are also pr
38、edicted reason-ably well over the investigated Math number range from 0.6 to 2.0 and at angles of attackfrom 0 to 60 . As for the bodies of revolution, the predictions are best for supersonic Mathnumbers.For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4,measured nor
39、mal-force coefficients and centers are predicted reasonably well at the uppertest Math number of 2.0. However, with a decrease in Math number to Moo = 0.6, theagreement for CN rapidly deteriorates, although the normal-force centers remain in closeagree me n t.For Moo=0.6, 0.9, and 2.0 and angles of
40、attack of 10 , 20 , 30 , 40 , and 50,vapor-screen and oil-flow pictures are shown for many body, body-wing, and body-wing-tailconfigurations. When separation and vortex patterns are asymmetric, undesirable side forcesare measured for the models even at zero sideshp angle.These side forces can be sig
41、nificantly affected by changes ir_ Mach number, nosefineness ratio, nose bluntness, and body cross section. Generally, the side-force coefficientsdecrease or vanish with the following: increase in Math number, decrease in nose finenessratio, change from sharp to blunt nose, and flattening of body cr
42、oss section iparticularly thebody nose). Additions of afterbody strakes, wings, or wings plus tail produce much smalleror no appreciable effects.*Formerly issued as NASA TM X-73,123 with limited distribution. Reissued for unlimited distribution.Provided by IHSNot for ResaleNo reproduction or network
43、ing permitted without license from IHS-,-,-CHAPTER 1INTRODUCTIONOver the last several years, high angle-of-attack aerodynamics has increased in impor-tance because of the demand for greater maneuverability of space shuttle vehicles, missiles,and military aircraft (both manned and remotely piloted).
44、Until recently there has been ageneral lack of analytical methods and aerodynamic data suitable for use in the preliminarydesign of most advanced configurations for flight to high angles of attack over a wide rangeof Mach and Reynolds numbers. There has been, however, considerable research leading t
45、othe development of methods for predicting the static aerodynamic characteristics of simpleshapes, primarily slender bodies of revolution.Prior to the work of Allen in 1949 (ref. 1), most analytical procedures for computingthe aerodynamic characteristics of bodies were based on potential-flow theory
46、, and theirusefulness was limited to very low angles of attack. Allen proposed a method for predictingthe static longitudinal forces and moments for bodies of revolution inclined to angles ofattack considerably higher than those for which theories based only on potential-flow con-cepts are known to
47、apply. In this method, a crossflow lift attributed to viscous-flow separa-tion is added to the lift predicted by potential theory. This method has been used quite suc-cessfully to compute the aerodynamic coefficients of inclined bodies (e.g., refs. 1-6),although most data available for study until 1
48、961 were for bodies at angles of attack belowabout 20 , and the formulas were initially written to apply only over about this angle-of-attack range.In 1961, Allens concept was adapted by Jorgensen and Treon (ref. 7) for computingthe normal-force, axial-force, and pitching-moment coefficients for a rocket boosterthroughout the angle-of-attack range from 0 to 180 . Reasonable agreement of theory withexperiment was obtained for a test model of the rocket booster over the Mach number cangefrom 0.6 to 4. The Allen concept
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