Showing papers in "Journal of Aircraft in 1978"

Design of Subsonic Airfoils for High Lift

Abstract: Nomenclature c = airfoil chord CL = lift coefficient = L/!/2pV00c CLu = upper-surface lift coefficient Cp = pressure coefficient = (p -p^)/ Ap Vx 2 Mx = freestream Mach number p = static pressure Re^ = freestream Reynolds number based on airfoil chord = V^clv sp = location of leading-edge stagnation point V^ — freestream velocity v local velocity on airfoil surface x = distance along chord line F = circulation about the airfoil 7 = ratio of specific heats v = kinematic viscosity p = density () oo = freestream conditions () t e = conditions at the airfoil trailing edge

Lift Enhancement by an Externally Trapped Vortex

Abstract: A theoretical study is made of the performance capabilities of a lift concept that utilizes a spanwise vortex over the upper surface of the wing. The vortex is generated by a vertical flap near the leading edge of the wing and maintained by suction through orifices in endplates at the wingtip. The analysis approximates the three-dimensional flow field with a two-dimensional configuration that is mapped by conformal transformation into the flow about a circle. Theoretical solutions for a range of flap and orifice configurations predict that section lift coefficients up to around 10 can be achieved. It is concluded that such a lift concept is applicable to STOL aircraft if the vortex can be adequately stabilized and if the endplate suction can be generated efficiently.

Helicopter noise: state-of-the-art

Abstract: This paper reviews helicopter external noise with particular emphasis on the noise due to helicopter main and tail rotors. The bases for annoyance and audibility are discussed. Sources of rotor noise include steady, periodic, and random loads on the rotor blades, as well as volume displacement and nonlinear aerodynamic effects at high blade Mach numbers. Either main or tail rotors can be dominant noise sources at various frequencies and observer positions.

Drones for aerodynamic and structural testing /DAST/ - A status report

Abstract: A program for providing research data on aerodynamic loads and active control systems on wings with supercritical airfoils in the transonic speed range is described. Analytical development, wind tunnel tests, and flight tests are included. A Firebee II target drone vehicle has been modified for use as a flight test facility. The program currently includes flight experiments on two aeroelastic research wings. The primary purpose of the first flight experiment is to demonstrate an active control system for flutter suppression on a transport-type wing. Design and fabrication of the wing are complete and after installing research instrumentation and the flutter suppression system, flight testing is expected to begin in early 1979. The experiment on the second research wing - a fuel-conservative transport type - is to demonstrate multiple active control systems including flutter suppression, maneuver load alleviation, gust load alleviation, and reduce static stability. Of special importance for this second experiment is the development and validation of integrated design methods which include the benefits of active controls in the structural design.

An Investigation of Strut-Wall Intersection Losses

Abstract: A study of the intersection losses associated with the junction of a symmetric airfoil and a planar wall is reported. An experimental program, conducted in a low-speed air tunnel, provided detailed wake total pressure profiles as well as surface flow visualization photographs which define the overall flowfield. The behavior of the intersection losses was examined for dependence on flow incidence angle and strut contour. It was found that the endwall intersection losses were strongly dependent on the thickness of the incident boundary layers-thick boundary layers producing markedly lower losses than very thin incident boundary layers. A heuristic model of the flowfield, which explains marked differences between the thick and thin boundary-layer results, is also presented.

STOE Potential of the Circulation Control Wing for High-Performance Aircraft

Abstract: Research and development being conducted at the David Taylor Naval Ship R and D Center investigates the STOL capability of the Circulation Control Wing (CCW) concept on high-performa nce aircraft. This high-lift system, which employs tangential blowing over a rounded trailing edge and requires mass flows characteristic of state-of-the-art turbine engine bleed, has demonstrated the ability to more than double the lift capability of conventional Navy and Marine aircraft. The resulting reduced takeoff and landing speeds and distances, plus increased overload capability, are achieved without severe compromise of wing structure, weight, or engine arrangement, and without large quantities of ducted hot gas. Based on these anticipated benefits and the results of existing experimental investigations, a program has been initiated to demonstrate the STOL capability of the CCW concept applied to a full-scale A-6 flight demonstrator aircraft. This paper will address the experimental development and optimization of the CCW system on an A-6 model and will present predicted full-scale STOL performance gains for the flight demonstrator.

Properties of Aspect-Ratio-4.0 Rectangular Jets in a Subsonic Crossflow

Abstract: A series of wind-tunnel and laboratory tests were conducted at the NASA Langley V/STOL tunnel facility to determine both the detailed structure and the induced effects of aspect-ratio-4.0 rectangular jets both in a subsonic crosswind and in quiescent conditions. Wind-tunnel tests were conducted on both blunt (nozzle major axis normal to free stream) and streamwise (nozzle major axis parallel to free stream) nozzle orientations for jet injection angles ranging from 15 to 90 degrees at jet-to-crossflow velocity ratios of 4, 8, and 10. Results indicate that the blunt nozzle induced effects are more significant than those produced by comparable streamwise-oriented jets and that both the flow-field structure and induced effects of streamwise-oriented rectangular jets are quite similar to those created by round jets. Additionally, it is shown that significant differences exist in the vortex flow fields generated by the same rectangular nozzle mounted in two different test hardware configurations.

Turbulent Airfoils for General Aviation

Abstract: Several new airfoils are presented which have short pieces of steep favorable pressure gradient followed by an early pressure recovery which is a compromise between the Stratford distribution and soft stall. The drag polars are computed by a mathematical model, which is briefly described. For comparison, NACA 652 - 415 and NASA GA(W)-1 airfoils are evaluated using the same model; in this case the results are also compared with experimental drag polars.

Effect of wing fins on lift-generated wakes

Abstract: A theoretical and experimental study has been made of the effect of wing-mounted fins on the vortex wakes of subsonic aircraft. The lateral lift on the fins injects vortices into the wake and redistributes the lift on the wing. The revised wake vorticity then interacts convectively to form a new configuration with low rotational velocities. The theory is used 1) to gain an understanding of wake alleviation by vortex injection and 2) to guide the experimental investigation. Wind-tunnel tests were used to evaluate the alleviation achievable and to find the optimum values for the various fin parameters. It was found that vertical fins mounted on the upper surface of a wing could lower the wake-induced rolling moments on an encountering wing by a factor of 3 or more. The most promising fin configuration found for the Boeing 747 model is a fin positioned 48% outboard from the centerline to the wingtip with a height equal to 0.014 wingspan, a chord equal to 0.085 wingspan, and an 18-deg angle of attack. This fin configuration caused a 10% increase in drag but no lift penalty.

Investigation of Interior Noise in a Twin-Engine Light Aircraft

Abstract: This paper describes experimental studies of interior noise in a twin-engine, propeller-driven, light aircraft. An analytical model for this type of aircraft is also discussed. Results indicate that interior noise levels in this aircraft due to propeller noise can be reduced by reducing engine rpm at constant airspeed (about 3 dB), and by synchrophasing the twin engines/propellers (perhaps up to 12 dB). Ground tests show that the exterior noise pressure imposed on the fuselage consists of a complex combination of narrow-band harmonics due to propeller and engine exhaust sources. This noise is reduced by about 20-40 dB (depending on the frequency) by transmission through the sidewall to the cabin interior. The analytical model described uses modal methods and incorporates the flat-side geometrical and skin-stringer structural features of this light aircraft.

Quantification of Fastener- Hole Quality

Abstract: This paper describes a method, the Equivalent Initial Quality Method, of quantifying the quality of fastener holes. This quantification is accomplished by representing the imperfections that are either inherent in a material or introduced during the manufacturing of a structural component with a fatigue crack of a particular size and shape. This initial quality representation can be used in a crack-propagation analysis to determine the life of the structural component. For example, the Equivalent Initial Quality Method can be used in design by providing the assumptions necessary to satisfy the USAF airplane damage tolerance design requirements (MIL-A-83444) and the USAF airplane durability design requirements (MIL-A-8866B). The method was used to assess the manufacturing quality of the F-4C/D/E(S) and A-7D aircraft. The more recent A-7D quality assessment is discussed in detail and the equivalent initial quality results for the F-4C/D and A-7D aircraft are presented. The potential applications (e.g., determination of required inspection intervals and maintenance and modification schedules, use in design, assessment of quality of manufacturing procedures, etc.), as well as the possible limitations (e.g., sensitivity of method to stress level, material, etc.), of the Equivalent Initial Quality Method are discussed.

Surface pressure measurements at the tip of a model helicopter rotor in hover

Abstract: Surface pressures were measured near the tip of a hovering single-bladed model helicopter rotor with two tip shapes The rotor had a constant-chord, untwisted blade with a square, flat tip which could be modified to a body-of-revolution tip Pressure measurements were made on the blade surface along the chordwise direction at six radial stations outboard of the 94 percent blade radius Data for each blade tip configuration were taken at blade collective pitch angles of 0, 618 and 114 degrees at a Reynolds number of 736,000 and a Mach number of 025 both based on tip speed Chordwise pressure distributions and constant surface pressure contours are presented and discussed

Positive Tail Loads for Minimum Induced Drag of Subsonic Aircraft

Abstract: By applying Prandtl's relation for the induced drag of a biplane to typical wing-tail combinations, it can be shown that the minimum induced drag occurs with a positive tail upload. This fact has been overlooked because the reduction in the total induced drag by a tail download was overestimated by using the total downwash of the wing on the tail, while neglecting the downwash produced on the wing by the tail. It is proved that, regardless of the relative size of the tail, the downwash produced by a tail download increases the induced drag of the wing so as to cancel the additional "tail thrust," and keep the mutually induced drag of a wing-tail combination the same as that induced upon the tail alone when it is in the wing's Trefftz-plane. At any finite tail length the bound circulation vortex of the wing produces a downwash that increases the induced drag of a tail upload. However, the circulation vortex system of the tail upload produces an upwash on the wing that results in a "wing thrust" component that cancels the increased drag on the tail so that the total induced drag is a minimum with a positive tail load. In order to facilitate the calculation of the mutually induced drag of typical wing-tail combinations, an explicit relation is derived for the limiting case of a small-span tail at any distance above or below a large-span wing.

Numerical modeling of three-dimensional flows in turbofan engine exhaust nozzles

Abstract: An experimental and numerical study was conducted to investigate the use of three-dimensi onal viscous analysis for the design of internal mixers for jet noise suppression. Three flows, a full-scale free mixer, a fullscale lobed mixer, and a model-scale lobed mixer, were selected for study. Details of the experiments, turbulence modeling, and the determination of initial flow properties from engine operating properties are described together with a discussion of the usefulness of the analysis for mixing design. Overall the results are encouraging and, at least for the flows considered here, numerical analysis appears to compare favorably with model-scale tests as a simulation of the full-scale flow.

A Basic Study of the VTOL Ground Effect Problem for Planar Flow

Abstract: The force interaction between airframe undersurfaces and the ground in the presence of lift jets is an important consideration for VTOL aircraft design. As a first step toward prediction of this phenomenon, a combined theoretical and experimental investigation of planar turbulent jet impingement flowfields has been undertaken. Unvectored jets in close ground effect have been modeled using the incompressible Reynolds equations with a one-equation turbulence model. Distributions of the flow properties are computed as functions of undersurface shape, length scales, and jet exit height above ground. Computed flowfield properties are presented and comparisons are made with experimental measurements.

Influence of Compressor Exit Conditions on Diffuser Performance

Abstract: Results are presented of an experimental investigation into the way in which compressor exit conditions influence the performance of two optimum combustor-dump diffusers. In an optimum system nearly all of the pressure rise occurs in a relatively long prediffuser which is designed to produce a symmetrical outlet velocity profile at the design flow split. One diffuser was tested downstream of a seven-stage axial flow compressor and, for comparative purposes, retested with fully developed flow at inlet. The second diffuser was tested downstream of an annular tandem cascade which could be sited at a number of positions relative to diffuser inlet. The overall performance, measured when the wakes from the outlet guide vanes had decayed considerably, was the same as that achieved with fully developed inflow.

Subsonic Vortex-Flow Design Study for Slender Wings

Abstract: A theoretical study describing the effects of spanwise camber on the lift dependent drag of slender delta wings having leading-edge vortex flow is presented. The earlier work by Barsby, using conical flow, indicated that drag levels similar to those in attached flow could be obtained. This is re-examined and then extended to the more practical case of nonconical flow by application of the vortex-lattice method coupled with the suction analogy and the recently developed Boeing free-vortex-sheet method. Lastly, a design code is introduced which employs the suction analogy in an attempt to define "optimum" camber surfaces for minimum lift dependent drag for vortex flow conditions.

Some observations on the mechanism of aircraft wing rock

Abstract: A pressure scale model of Northrop F-5A was tested in NASA Ames Research Center Eleven-Foot Transonic Tunnel to simulate the wing rock oscillations in a transonic maneuver. For this purpose, a flexible model support device was designed and fabricated which allowed the model to oscillate in roll at the scaled wing rock frequency. Two tunnel entries were performed to acquire the pressure (steady state and fluctuating) and response data when the model was held fixed and when it was excited by flow to oscillate in roll. Based on these data, a limit cycle mechanism was identified which supplied energy to the aircraft model and caused the Dutch roll type oscillations, commonly called wing rock. The major origin of the fluctuating pressures which contributed to the limit cycle was traced to the wing surface leading edge stall and the subsequent lift recovery. For typical wing rock oscillations, the energy balance between the pressure work input and the energy consumed by the model aerodynamic and mechanical damping was formulated and numerical data presented.

Wing Planforms for Large Military Transports

Abstract: Transport aircraft, designed for long-range military missions with heavy payloads, lead to wings with high aspect ratios and very large spans. A wing-geometry/cruise-speed optimization study was made of a large cantilever-wing military transport airplane. Preliminary design and performance evaluations were also made of a strut-braced wing airplane. Initial results obtained with statistical weights indicated small performance advantages for the cantilever-wing design. Subsequent results obtained with weights derived from detailed analytical structural analyses reversed the initial conclusions. These results indicated that unusual alternative configuration concepts cannot be discarded, based on small differences predicted during conceptual design studies.

Wing Store Active Flutter Suppression-Correlation of Analyses and Wind-Tunnel Data

Abstract: Through analyses and recent wind-tunnel tests, active flutter suppression has been shown to be a promising technique for preventing wing/external store flutter restrictions. Data measured in the wind tunnel have been used to evaluate the validity of a method for the design and analysis of active feedback control systems. The results of an analytical effort to study the behavior of an active flutter suppression wind-tunnel model are presented and compared with available test data. For this application, the model was aerodynamically represented by subsonic doublet lattice theory and stability was evaluated using modified Nyquist criteria.

Minimum-Time Loop Maneuvers of Jet Aircraft

Abstract: Minimum-time loop maneuvers of jet aircraft are investigatedvby means of the calculus of variations. The optimal control (lift coefficient and thrust) is determined as a function of the state variables and Lagrange multipliers for an arbitrary maneuver in the vertical plane. Intermediate thrust arcs and intermediate lift arcs with minimum thrust are shown to be nonoptimal for the dynamic models considered. Rules are established for joining trajectory segments with different types of control behavior. Using these results, optimal loop trajectories are generated for a given set of initial conditions and a large number of values of maximum lift coefficient and maximum thrust. The control history and the maneuver time are found to be strongly affected by the latter two parameters. Two aerodynamic models are considered, one relatively simple and the other somewhat more complicated.

Transonic Shock-Turbulent Boundary Layer Interaction with Suction and Blowing.

Abstract: : A basic theory of weak normal shock - turbulent boundary layer interactions is given for two-dimensional non-separating flows including mass transfer across the wall throughout the interaction region. Even small amounts of suction are found to significantly reduce both the streamwise scale and thickening effect of the interaction and delay the onset of separation. This is shown to be a consequence of the large mass transfer effect on the shape of the incoming boundary layer Mach number profile away from the wall. Parametric study results showing the influence of Reynolds and shock Mach number as well as mass transfer parameter on the interaction, plus favorable comparisons with various experimental data, are also presented.

Noise Due to Interaction of Boundary-Layer Turbulence with a Compressor Rotor

Abstract: The radiated sound due to a compressor or propulsor rotating blade row was investigated under various operating conditions and inflows. The propulsor was operated in air with different blade space-to-chord ratios, different flow coefficients and differing turbulence (nonisotropic) inflows. The inflows ingested were: 1) the natural boundary layer on the hub and annulus wall, 2) a tripped boundary layer on the hub, and 3) a fully developed boundary layer on the hub. The turbulence properties were also altered by placing a grid at the inlet. The mean velocity profiles, turbulence intensities, length scales, and energy spectra of the inflow, as well as nearand far-field acoustic spectra, were measured. A parametric investigation of the effect of inflow characteristics on the radiated sound was made. Several length scales were found to exist simultaneously. The noise due to small scale turbulence seemed to depend on the ratio of the square of the turbulence velocity normal to the blade divided by the axial length scale. The long eddies (compared to blade spacing) were primarily responsible for discrete tone production. Nomenclature ABL,NBL,FDBL = artificial, natural, and fully developed boundary layer, respectively B = number of blades c = blade chord CL =lift coefficient based on the cascade mean velocity Eu.u.(k) = spectrum function of turbulence kinetic energy k = wave number

Detecting abnormal turbine engine deterioration using electrostatic methods

Abstract: A method of detecting abnormal turbine engine deterioration has been developed and tested. The method observes pulse electrostatic signals in the exhaust which have been determined to originate from component rubbing, chaffing, erosion, and burning (i.e., various forms of deterioration). The normal (healthy engine) deterioration rate is first studied as a function of engine cycling and power. This deterioration rate is then normalized with an engine power and an engine cycling parameter. Tenfold increases in the normalized deterioration rate are then used as an indication of impending component failure. Experience shows that about two out of three turbine engines gas-path failures can be predicted four or more hours ahead of time by this method. The false alarm rate is estimated to be about 5%.

A Potential Flow Design Method for Multicomponent Airfoil Sections

Abstract: A multicomponent, potential flow design method has been developed to generate an airfoil section with a specified velocity distribution on its surface The method was developed from an accurate and efficient surface vorticity analysis method An iterative approach is used to adjust the geometry of a basic airfoil section until it gives the required velocity distribution This approach allows the designer to adjust the solution to suit the practical constraints within which airfoil designers must work Examples of the results of various airfoil design problems are shown

Parachute Canopy Stress Measurements at Steady State and During Inflation

Abstract: Circumferential stresses were measured on 5-ft flat circular block-constructed solid cloth and ringslot parachutes in a wind tunnel during inflation under infinite mass condition and at steady state. Five Omega senors placed along gore centerlines and a force link measured stresses and the parachute force versus time. Radial stresses were measured on a 3-ft solid flat circular block-constructed parachute at steady state. Stress measurements were evaluated in view of maximum and steady-state stresses versus location and time, as well as in respect to the parachute force. The measurements indicated strong stress variations, being different for both types of parachutes. In the 3-ft solid cloth parachute, the radial stress was, at certain locations, greater than the circumferential stress.