Engineering topics

Bingham, G. J.

Publications and source records attributed to Bingham, G. J..

12 records

NASA NTRS1985

Optimization methods applied to the aerodynamic design of helicopter rotor blades

This paper describes a formal optimization procedure for helicopter rotor blade designs which minimizes hover horsepower while assuring satisfactory forward flight performance. The approach is to couple hover and forward flight analysis programs with a general purpose optimization procedure. The resulting optimization system provides a systematic evaluation of the rotor blade design variables and their interaction, thus reducing the time and cost of designing advanced rotor blades. The paper discusses the basis for and details of the overall procedure, describes the generation of advanced blade designs for representative Army helicopters, and compares designs and design effort with those from the conventional approach which is based on parametric studies and extensive cross-plots.

Walsh, J. L.↗

NASA NTRS1984

Shapes for rotating airfoils

An airfoil which has particular application to the blade or blades of rotor aircraft and aircraft propellers is presented. The airfoil thickness distribution, camber and leading edge radius are shaped to locate the airfoil crest at a more aft position along the chord, and to increase the freestream Mach number at which sonic flow is attained at the airfoil crest. The reduced slope of the airfoil causes a reduction in velocity at the airfoil crest at lift coefficients from zero to the maximum lift coefficient. The leading edge radius is adjusted so that the maximum local Mach number at 1.25 percent chord and at the designed maximum lift coefficient is limited to about 0.48 when the Mach number normal to the leading edge is approximately 0.20. The lower surface leading edge radius is shaped so that the maximum local Mach number at the leading edge is limited to about 0.29 when the Mach number normal to the leading edge is approximately 0.20. The drag divergence Mach number associated with the airfoil is moved to a higher Mach number over a range of lift coefficients resulting in superior aircraft performance.

Bingham, G. J.↗

NASA NTRS1984

Helicopter rotor blade aerodynamic optimization by mathematical programming

Formal mathematical programing was applied to the aerodynamic rotor blade design process. The approach is to couple hover and forward flight analysis programs with the general-purpose optimization program CONMIN to determine the blade taper ratio, percent taper, twist distribution, and solidity which minimize the horsepower required at hover while meeting constraints on forward flight performance. Designs obtained using this approach for the blade of a representative Army helicopter compare well with those obtained using a conventional approach involving personnel-intensive parametric studies. Results from the present method can be obtained in 2 days as compared to 5 weeks required by the conventional procedure. Also the systematic manipulation of the design variables by the optimization procedure minimizes the need for the researcher to have a vast body of past experience and data in determining the influence of a design change on the performance.

Walsh, J. L.↗

NASA NTRS1982

Two-dimensional aerodynamic characteristics of three rotorcraft airfoils at Mach numbers from 0.35 to 0.90

Three airfoils designed for helicopter rotor application were investigated in the Langley 6- by 28-inch Transonic Tunnel to determine the two dimensional aerodynamic characteristics at Mach numbers from 0.34 to 0.88 and respective Reynolds numbers from about 4.4 x 10(6) power to 9.5 x 10(6) power. The airfoils have thickness-to-chord ratios of 0.08, 0.10, and 0.12. Trailing-edge reflex was applied to minimize pitching moment. The maximum normal-force coefficient of the RC(3)-12 airfoil is from 0.1 to 0.2 higher, depending on Mach number M, than that of the NACA 0012 airfoil tested in the same facility. The maximum normal-force coefficient of the RC(3)-10 is about equal to that of the NACA 0012 at Mach numbers to 0.40 and is higher than that of the NACA 0012 at Mach numbers above 0.40. The maximum normal force coefficient of the RC(3)-08 is about 0.19 lower than that of the NACA 0012 at a Mach number of 0.35 and about 0.05 lower at a Mach number of 0.54. The drag divergence Mach number of the RC(3)-08 airfoil at normal-force coefficients below 0.1 was indicated to be greater than the maximum test Mach number of 0.88. At zero lift, the drag-divergence Mach numbers of the RC(3)-12 and the RC(3)-10 are about 0.77 and 0.82, respectively.

Bingham, G. J.↗

NASA NTRS1981

Two-dimensional aerodynamic characteristics of an airfoil designed for two

An airfoil designed for helicopter rotor application is investigated. The airfoil is designed to increase maximum normal force coefficient while maintaining favorable drag divergence and pitching moment characteristics. Two modifications are also tested. Maximum normal force coefficient varies from 1.14 to 0.90 at Mach numbers from about 0.35 to 0.65. Both modifications decreased drag coefficient at zero normal force coefficient for Mach numbers near drag divergence, but were less beneficial at a normal force coefficient of -0.2.

Bingham, G. J.↗

NASA NTRS1980

Aerodynamic characteristics of three helicopter rotor airfoil sections at Reynolds number from model scale to full scale at Mach numbers from 0.35 to 0.90

An investigation was conducted in the Langely 6 by 28 inch transonic tunnel to determine the two dimensional aerodynamic characteristics of three helicopter rotor airfoils at Reynolds numbers from typical model scale to full scale at Mach numbers from about 0.35 to 0.90. The model scale Reynolds numbers ranged from about 700,00 to 1,500,000 and the full scale Reynolds numbers ranged from about 3,000,000 to 6,600,000. The airfoils tested were the NACA 0012 (0 deg Tab), the SC 1095 R8, and the SC 1095. Both the SC 1095 and the SC 1095 R8 airfoils had trailing edge tabs. The results of this investigation indicate that Reynolds number effects can be significant on the maximum normal force coefficient and all drag related parameters; namely, drag at zero normal force, maximum normal force drag ratio, and drag divergence Mach number. The increments in these parameters at a given Mach number owing to the model scale to full scale Reynolds number change are different for each of the airfoils.

Noonan, K. W.↗

NASA NTRS1979

Experimental investigation of three helicopter rotor airfoils designed analytically

Three helicopter rotor airfoils designed analytically were investigated in a wind tunnel at Mach numbers from about 0.30 to 0.90 and Reynolds from about 0.8 to 2.3 x 10 to the 6th power. The airfoils had thickness-to-chord ratios of 0.08, 0.10, and 0.12 with maximum thickness at 40 percent chord. The camber distribution of each section was the same with maximum camber at 35 percent chord. The 10-percent-thick airfoil was also investigated at Reynolds numbers from 4.8 to 9.4 x 10 to the 6th power. The drag divergence Mach number of the 10-percent-thick airfoil is about 0.83 at a normal-force coefficient of 0 and about 0.72 at a normal-force coefficient of 0.6 at Reynolds numbers near 9 x 10 to the 6th power. The maximum normal-force coefficient is slightly less than that of the NACA 0012 airfoil tested in the same facility. The results indicate that a qualitative evaluation of the drag divergence can be made at normal-force coefficients up to the onset of boundary-layer separation by analytically predicting the onset of sonic flow at the airfoil crest. The qualitative results are conservative with respect to experimental values with the experimental drag divergence Mach number up to 0.05 higher than that indicated by analysis.

Bingham, G. J.↗

NASA NTRS1977

Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

Noonan, K. W.↗

NASA NTRS1975

An analytical evaluation of airfoil sections for helicopter rotor applications

An analytical technique was used to evaluate airfoils for helicopter rotor application. This technique permits assessment of the influences of airfoil geometric variations on drag divergence Mach number at lift coefficients from near zero to near maximum lift. Analytical results presented in this paper indicate the compromises in drag divergence Mach number which result from changes in (1) thickness ratio, (2) location of maximum thickness, (3) leading-edge radius, (4) camber addition, and (5) location of maximum camber of NACA four- and five-digit-series airfoils and some 6-series airfoils of potential interest for helicopters. Examples of airfoil sections which combine several of the geometric changes favorable to both advancing and retreating section performance have been presented.

Bingham, G. J.↗

NASA NTRS1975

An analytical and experimental evaluation of airfoil sections for helicopter rotor application

The influence of the more independent airfoil parameters such as thickness, thickness distribution, leading-edge radius, camber, and camber distribution on lift-Mach number characteristics is investigated at lift coefficients up to near-maximum lift. The analysis is based on the drag divergence Mach number (Md) prediction techniques, where Md is the free-stream Mach number at which the rate of increase of drag coefficient with Mach number equals 0.1. The analytical results obtained indicate the compromises in Md which result from changes in thickness ratio, location of maximum thickness, leading edge radius, camber addition, and location of maximum camber for four- and five-digit airfoils and some six-series airfoils of potential interest for helicopters. An example of airfoil sections which combines several of the favorable geometric changes is evaluated analytically and experimentally. A comparison of results shows that the relative effect of the geometric changes on the lift coefficient-Md relation is realistic, and that the methods of analysis employed can be effectively used during preliminary vehicle design and airfoil selection.

Bingham, G. J.↗

NASA NTRS1974

Low speed aerodynamic characteristics of NACA 6716 and NACA 4416 airfoils with 35 percent-chord single-slotted flaps

An investigation was conducted in a low-turbulence pressure tunnel to determine the two-dimensional lift and pitching-moment characteristics of an NACA 6716 and an NACA 4416 airfoil with 35-percent-chord single-slotted flaps. Both models were tested with flaps deflected from 0 deg to 45 deg, at angles of attack from minus 6 deg to several degrees past stall, at Reynolds numbers from 3.0 million to 13.8 million, and primarily at a Mach number of 0.23. Tests were also made to determine the effect of several slot entry shapes on performance.

Bingham, G. J.↗

NASA NTRS1972

Low-speed aerodynamic characteristics of an airfoil optimized for maximum lift coefficient

An investigation has been conducted in the Langley low-turbulence pressure tunnel to determine the two-dimensional characteristics of an airfoil optimized for maximum lift coefficient. The design maximum lift coefficient was 2.1 at a Reynolds number of 9.7 million. The airfoil with a smooth surface and with surface roughness was tested at angles of attack from 6 deg to 26 deg, Reynolds numbers (based on airfoil chord) from 2.0 million to 12.9 million, and Mach numbers from 0.10 to 0.35. The experimental results are compared with values predicted by theory. The experimental pressure distributions observed at angles of attack up to at least 12 deg were similar to the theoretical values except for a slight increase in the experimental upper-surface pressure coefficients forward of 26 percent chord and a more severe gradient just behind the minimum-pressure-coefficient location. The maximum lift coefficients were measured with the model surface smooth and, depending on test conditions, varied from 1.5 to 1.6 whereas the design value was 2.1.

Bingham, G. J.↗