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At least 235 records · Page 13

All sources NASA NTRS DOE OSTI

NASA NTRSDate not supplied

Results from Two Full-Scale Hawker 4000 Fuselage Drop Tests

In the spring of 2022, NASA Langley Research Center (LaRC) conducted two full-scale drop tests on a Hawker 4000 partial fuselage structure for the evaluation of composite material response under dynamic impact loading conditions. The specific objectives of the tests were to induce dynamic composite damage initiation and progression and in primary structure and to understand the nature of the failure modes. The tests were conducted via dividing the fuselage hardware into three smaller sections. Dynamic drop tests were conducted on the forward and aft portions of the fuselage under different dynamic loading conditions, while the middle section was reserved for materials testing. Both tests were conducted by dropping the test articles from a height of 14 feet with significant amounts of pitch, which was added to initiate damage in specific regions of the structure. The test articles were ballasted then instrumented with accelerometers at various locations and speckle-coated on both exterior sides with a black and white stochastic pattern for use with three-dimensional digital image correlation (3D-DIC). In addition, the interior belly portion of the forward section test article was also speckle-coated, which was intended to capture deformations on the belly of the test article from the interior at impact. The interior of the aft section test article was not speckle-coated. Prior to testing, bow-tie markers were applied at various locations for use in marker tracking, to measure impact conditions such as vertical impact velocity and pitch. In this report, a partial summary of the data collected from the tests is presented. The data was primarily in the form of accelerometers to measure impact acceleration and 3D-DIC to measure deformation and failure onset and propagation. Post-test inspections revealed the primary damage mechanism was fiber cracking and delamination primarily in and around the area that contacted the impact surface. Acceleration pulse shapesin the ballast locations were trapezoidal in nature and ranged between 10 g and 40 g, depending on the location and test.

full scale testing↗

NASA NTRSDate not supplied

Results from Two Full-Scale Hawker 4000 Fuselage Drop Tests

full scale testing↗

NASA NTRS1952

Investigation of the Influence of Fuselage and Tail Surfaces on Low-speed Static Stability and Rolling Characteristics of a Swept-wing Model

Results are presented of a wind-tunnel investigation to determine influence of the fuselage and tail on static stability and rotary derivatives in roll of a model having 45 degrees sweptback wing and tail surfaces. The wing alone and the model without the horizontal tail showed marginal longitudinal stability near maximum lift. The longitudinal stability of the complete model was satisfactory. The vertical tail produced larger increments of rate of change of lateral-force and yawing-moment coefficients with wing-tip helix angle than the fuselage or the horizontal tail.

TAIL-WING-FUSELAGE COMBINATIONS - AIRPLANES↗

NASA NTRS1957

Aerodynamic Forces and Moments on a Large Ogive-Cylinder Store at Various Locations Below the Fuselage Center Line of a Swept-Wing Bomber Configuration at a Mach Number of 1.61

A supersonic wind-tunnel investigation on store interference has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 1.61. Forces and moments were measured on a large ogive-cylinder store in the presence of a 45 deg swept-wing-fuselage bomber configuration for a number of store locations below the fuselage center line. Results of the investigation show that large variations of store lift, drag, and pitch occur with changes in store or airplane angle of attack, store vertical location, and store horizontal location. The variation of the store forces and moments with respect to the chordwise location of the wing plan form indicates that the wing is a large factor in producing the interference loads on the store. Comparison of data for underfuselage and underwing store locations at an angle of attack of 0 deg showed maximum store drag interferences of similar magnitudes, but showed considerably smaller maximum interference on store lift an pitching moments for underfuselage store locations.

SUPERSONIC FLIGHT↗

NASA NTRS1957

Interference Effects of Fuselage-Stored Missiles on Inlet Duct Model of an Interceptor-Type Aircraft at Mach Numbers 1.5 to 1.9

The effect of missile armament on the performance of an interceptor-type aircraft model has been determined at Mach numbers 1.5, 1.7, and 1.9 and at angles of attack to 19 deg. With this configuration missiles were carried in a bay located on the bottom of the aircraft fuselage and mounted to a rotatable missile door. Rotation of the door then brought the missiles into the external or firing position. The aircraft model was characterized by triangular-shaped normal-shock inlets located at the wing roots. Relatively short and curved subsonic diffusers fed simulated twin side-by-side turbojet engines. Inasmuch as the missile bay extended considerably ahead of the inlet station, rotation of the missile door created considerable disturbance of the flow entering the inlets. In comparison with the internal missile arrangement, the external missile configurations increased the model lift, drag, and pitching moment. While the diffuser-exit flow distortion and static-pressure fluctuations were not greatly affected, diffuser total-pressure recovery was reduced as much as 0.058 at Mach number 1.9 for one missile configuration. The most detrimental effect of missile-door rotation occurred at the transient door positions, or with the door halfwzy between the missiles-in and -out conditions. At this door position the flow into the inlets was highly asymmetrical. Although the performance of both left and right ducts was generally reduced, the inlet duct on the cavity side of the missile door was most severely penalized, becoming unstable recovery losses and increases in flow distortion. The installation of fuselage fences along the missile bay was only partially effective in reducing these losses.

Piercy, T. G.↗

NASA NTRS1972

Application study of filamentary composites in a commercial jet aircraft fuselage

A study of applications of filamentary composite materials to aircraft fuselage structure was performed. General design criteria were established and material studies conducted using the 727-200 forebody as the primary structural component. Three design approaches to the use of composites were investigated: uniaxial reinforcement of metal structure, uniaxial and biaxial reinforcement of metal structure, and an all-composite design. Materials application studies for all three concepts were conducted on fuselage shell panels, keel beam, floor beams, floor panels, body frames, fail-safe straps, and window frames. Cost benefit studies were conducted and developmental program costs estimated. On the basis of weight savings, cost effectiveness, developmental program costs, and potential for early application on commercial aircraft, the unaxial design is recommended for a 5-year flight service evaluation program.

Johnson, R. W.↗

NASA NTRS1972

A mixed optimization method for automated design of fuselage structures.

A procedure for automating the design of transport aircraft fuselage structures has been developed and implemented in the form of an operational program. The structure is designed in two stages. First, an overall distribution of structural material is obtained by means of optimality criteria to meet strength and displacement constraints. Subsequently, the detailed design of selected rings and panels consisting of skin and stringers is performed by mathematical optimization accounting for a set of realistic design constraints. The practicality and computer efficiency of the procedure is demonstrated on cylindrical and area-ruled large transport fuselages.

Sobieszczanski, J.↗

NASA NTRS1973

Low-Speed Aerodynamic Characteristics of a Fuselage Model with Various Arrangements of Elongated Lift Jets

Data were obtained for a round jet located on the center of the bottom of a fuselage and for elongated slots separated spanwise by distances of 0.8 and 1.2 of the fuselage width. The effect of yawing the slots, inclining the jets laterally, and combining slot yaw with jet inclination was determined. Data were obtained in and out of ground effect through a range of effective velocity ratios and through a range of sideslip angles.

Vogler, R. D.↗

NASA NTRS1974

Flight-test investigation of the aerodynamic characteristics and flow interference effects about the aft fuselage of the F-111A airplane

Static pressure measurements were made on the aft fuselage of an F-111A airplane to determine local flow characteristics and engine/airframe interaction effects. Data were obtained over the Mach number range from 0.5 to 2.0. Aspiration effects associated with low ejector nozzle expansion ratios reduced the local pressure coefficients particularly on the interfairing but also extending to the trailing edge of the nacelle. The presence of afterbodies also affected the behavior of the air flowing into and about the ejector nozzle. Pressures about the aft fuselage were improved by an increase in primary nozzle area at a supersonic speed. A comparison of wind-tunnel and flight-test results showed generally good agreement, although there was a large disparity in pressure level about the ejector nozzle. However, the shape of the data curves and the local flow behavior were basically similar.

Taillon, N. V.↗

NASA NTRS1974

Lateral-directional stability characteristics of a wing-fuselage configuration at angles of attack up to 44 deg

An investigation has been conducted to determine the effects of configuration variables on the lateral-directional stability characteristics of a wing-fuselage configuration. The variables under study included variations in the location of a single center-line vertical tail and twin vertical tails, wing height, fuselage strakes, and horizontal tails. The study was conducted in the Langley high-speed 7-by 10-foot tunnel at a Mach number of 0.30, at angles of attack up to 44 deg and at sideslip angles of 0 deg and plus or minus 5 deg.

Henderson, W. P.↗

NASA NTRS1974

Wind tunnel investigation of helicopter rotor wake effects on three helicopter fuselage models

The effects of rotor downwash on helicopter fuselage aerodynamic characteristics were investigated. A rotor model for generating the downwash was mounted close to each of three fuselage models. The main report presents the force and moment data in both graphical and tabular form and the pressure data in graphical form. This supplement presents the pressure data in tabular form. Each run or parameter sweep is identified by a unique run number. The data points in each run are identified by a point number. The pressure data can be matched to the force data by matching the run and point number.

Wilson, J. C.↗

NASA NTRS1975

Preservation of wing leading edge suction at the plane of symmetry as a factor in wing-fuselage design

Most fuselage geometries cover a portion of the wing leading edge near the plane of symmetry, and it seems reasonable to expect that a large fraction of the leading edge suction which would be developed by the covered wing at high angles of attack is not developed on the fuselage. This is one of the reasons that the Oswald span efficiency factor for the wing body combination fails to approach the value predicted by lifting line theory for the isolated wing. Some traditional and recent literature on wing-body interference is discussed and high Reynolds number data on wing-body-nacelle drag are reviewed. An exposed central leading edge geometry has been developed for a sailplane configuration. Low Reynolds number tests have not validated the design concept.

Larrabee, E. E.↗

NASA NTRS1976

Effects of surface temperature and Reynolds number on heat transfer to the Shuttle Orbiter leeward fuselage

Heat-transfer data obtained at hypersonic shock tunnel conditions and three-dimensional flow field computations were used to study the influence of surface temperature and Reynolds number on the heating experienced by the leeward fuselage area of the Space Shuttle Orbiter configuration. The basic results of this study indicate that systematic variations in the average total enthalpy within a boundary layer (as obtained through controlled nonadiabatic processes) has an influence on the heat transfer to downstream areas which can be correlated; even in three-dimensional separated flow areas. Specifically, the average separated-flow Stanton number for the fuselage leeward surface is shown to be moderately dependent on the windward-wall to free-stream total temperature ratio.

Bertin, J. J.↗

NASA NTRS1976

A comparison of the aerodynamic characteristics at transonic speeds of four wing-fuselage configurations as determined from different test techniques, 4 October 1960

The high speed aerodynamic characteristics of a family of four wing-fuselage configurations of 0, 35, 45, and 60 deg sweepback were determined from transonic bump model tests that were conducted in the Langley high speed 7 by 10 foot tunnel; sting supported model tests were conducted in the Langley 8 foot high speed tunnel and in the Langley high speed 7 by 10 foot tunnel, and rocket model tests were conducted by the Langley Pilotless Aircraft Research Division. A complementary study of the effect of Mach number gradients and streamline curvature on bump results is also included. The qualitative data obtained from the various test facilities for the wing-fuselage configurations were in essential agreement as regards the relative effects of sweepback and Mach number except for drag at zero lift. Quantitatively, important differences were present.

Donlan, C. J.↗

NASA NTRS1978

A procedure for the determination of the effect of fuselage nose bluntness on the wave drag of supersonic cruise aircraft

The incremental wave drag penalty due to nose blunting of a fuselage was investigated using a three dimensional finite difference scheme. An aircraft typical of current supersonic cruise concepts was considered. Computational problems with the finite difference scheme as the fuselage afterbody closes were addressed. A linear theory method was employed to compute the afterbody aerodynamics and effectively extends the finite difference scheme to closing afterbodies. Acceptable drag increments for various levels of nose bluntness were demonstrated using this approach.

Walkley, K. B.↗

NASA NTRS1972

Tabulated pressure measurements of a NASA supercritical-wing research airplane model with and without fuselage area-rule additions at Mach 0.25 to 1.00

Basic pressure measurements were made on a 0.087-scale model of a supercritical wing research airplane in the Langley 8 foot transonic pressure tunnel at Mach numbers from 0.25 to 1.00 to determine the effects on the local aerodynamic loads over the wing and rear fuselage of area-rule additions to the sides of the fuselage. In addition, pressure measurements over the surface of the area-rule additions themselves were obtained at angles of sideslip of approximately - 5 deg, 0 deg, and 5 deg to aid in the structural design of the additions. Except for representative figures, results are presented in tabular form without analysis.

Harris, C. D.↗

NASA NTRS1978

A study of the structural-acoustic response and interior noise levels of fuselage structures

Models of both flat and curved fuselage panels were tested for their sound transmission characteristics. The effect of external air flow on transmission loss was simulated in a subsonic wind-tunnel. By numerically evaluating the known equations for field-incidence transmission loss of single-walled panels in a computer program, a comparison of the theory with the test results was made. As a further extension to aircraft fuselage simulation, equations for the field-incidence transmission loss of a double-walled panel were derived. Flow is shown to provide a small increase in transmission loss for a flat panel. Curvature is shown to increase transmission loss for low frequencies, while also providing a sharp decrease in transmission loss at the ring frequency of the cylindrical panel. The field-incidence transmission loss of a double-walled panel was found to be approximately twice that for a single-walled panel, with the addition of dips in the transmission loss at the air gap resonances and at the critical frequency of the internal panel.

Koval, L. R.↗

NASA NTRS1979

Recent progress in finite-volume calculations for wing-fuselage combinations

Progress in the application of finite-volume methods to the calculation of transonic potential flows past general wing-body combinations is reviewed. Two different methods of generating boundary-conforming grids are investigated, and the results compared to provide an estimate of solution sensitivity to grid geometry. Both conservative and quasi-conservative difference schemes are used in one of the coordinate systems. Results show that the error introduced by the quasi-conservative formulation seems to be small, although a one-dimensional analysis suggests that schemes of this type do not necessarily produce mass-conserving shocks. Comparison of calculated results with experimental data for realistic fuselage geometries clearly shows the importance of modelling the effect of fuselage geometry upon the wing pressure distribution.

Caughey, D. A.↗