{ "query": "transonic aileron buzz", "architecture": "arch-0.3-bge", "retrieval": "workers-ai-bge-query-embedding + opensearch-knn-hybrid", "embeddingModel": "@cf/baai/bge-base-en-v1.5", "vectorWeight": 0.7, "index": "cranfield-v0-bge-base-en-v15-gen023", "resultCount": 10, "latencyMs": 1247, "results": [ { "id": "496", "score": 1, "title": "a theory of transonic aileron buzz, neglecting viscous effects .", "abstract": "a theory of transonic aileron buzz, neglecting viscous effects . usaf-sponsored analysis of the unsteady perturbations of two-dimensional transonic flow around an airfoil, where local supersonic regions terminated by shock waves are present in the vicinity of the airfoil . viscous effects are neglected, and a linearized theory of the perturbations due to harmonic oscillations of an aileron is developed . a series solution for the pressure distribution is obtained, and numerical results for the nonsteady hinge moment, from the first approximation to the solution, are presented . as a result of flutter analysis a stability boundary for transonic aileron buzz is obtained . comparison of the theoretical results with experimental observations shows satisfactory agreement ." }, { "id": "903", "score": 0.3752, "title": "two dimensional transonic unsteady flow with shock waves .", "abstract": "two dimensional transonic unsteady flow with shock waves . a study is made of the unsteady flow around an airfoil at transonic mach numbers, the situation being such that local supersonic regions terminated by shock-waves are present in the vicinity of the airfoil . for the unsteady part of the flow, small perturbations technique is employed and the interaction with the shock wave is taken into account . the case of an oscillating aileron is considered first, and a solution is derived for the pressure distribution on the aileron . it is found that the solution has a simple form when the shock-wave is well ahead of the hinge axis of the aileron . as the shock approaches the hinge-axis a correction must be added to the solution . an interpretation of these results is given . the results are compared with results of a theory which neglects the presence of the shock and it is found that both agree for m = 1 . for m 1, however, neglecting the presence of the shock waves introduces errors of the order of magnitude (1 - m), where m is the local mach number behind the shock . the theory is finally extended to include the case in which the whole airfoil oscillates, but only the solution for the subsonic region behind the shock is treated . the role of the unsteady shock-boundary layer interaction is discussed and it is shown that this mechanism can be included in the results of the present theory ." }, { "id": "643", "score": 0.305, "title": "an investigation of wing-aileron flutter using ground launched rocket models .", "abstract": "an investigation of wing-aileron flutter using ground launched rocket models . control surface flutter of the wing torsion-control rotation type has been investigated for an unswept wing with an under-massbalanced, half span, outboard aileron . thirteen pairs of wings were tested, using ground launched rocket driven vehicles, and a range of values of aileron natural frequency was covered . the test results showed considerable scatter, but enabled upper and lower limits of a flutter boundary to be determined approximately . it was established that aileron flutter could be eliminated on the models tested provided the aileron frequency exceeded the wing torsional frequency by 20 per cent or more . in this condition the models were also free from single degree of freedom flutter" }, { "id": "199", "score": 0.2896, "title": "measurement of two dimensional derivatives on a wing-aileron-tab system .", "abstract": "measurement of two dimensional derivatives on a wing-aileron-tab system . measurements have been made of the direct two-dimensional damping and stiffness derivatives for a in incompressible flow . corrections arising from the apparatus are discussed and reference is made to an attempt to measure the direct tab derivatives . the effects are shown of frequency parameter, amplitude of oscillation, reynolds number, aileron angle and position of transition on the wing . variation with frequency parameter is substantially the same as for vortex-sheet theory and variation of amplitude produces little change in both derivatives . at the lowest reynolds number there is little change in both derivatives with variation of aileron angle for the condition of natural transition, but at higher reynolds numbers the stiffness derivatives increase at . a forward movement of transition reduces the stiffness derivatives at the smaller aileron angles, but at, at the lowest reynolds number, an increase results . similar trends are observed for the damping derivatives above . comparison with vortex-sheet theory shows that the measured values of the stiffness and damping derivatives are approximately 0.6 of the theoretical values . measurements have been made of the direct tab derivatives and cross aileron-tab derivatives for a per cent aileron and 4 per cent (approx.) tab . in addition some measurements of the direct aileron derivatives have been made for comparison with earlier results together with a number of static derivatives for the wing and controls . the influence is shown of frequency parameter, reynolds number, position of transition, mean tab angle and sealing of the control hinge gaps . some tests have been made with the ailcron set at minus 8 deg and the tab at plus 12 deg for which condition the hinge moment on the aileron was zero . reasonable agreement with the values given by the /equivalent profile/ theory is shown for both direct damping derivatives and for the direct tab stiffness derivative . the direct aileron stiffness derivative shows some departure from the theoretical value when . at and the natural transition, comparison with the values given by flat-plate theory gives the following approximate factors, where suffix denotes the theoretical values .." }, { "id": "1290", "score": 0.2039, "title": "measured and calculated subsonic and transonic flutter characteristics of a 45 sweptback wing planform in air and in freon-12 in the langley transonic dynamics tunnel .", "abstract": "measured and calculated subsonic and transonic flutter characteristics of a 45 sweptback wing planform in air and in freon-12 in the langley transonic dynamics tunnel . in order to investigate the reliability of flutter data measured in the langley transonic dynamics tunnel, an experimental and theoretical subsonic and transonic flutter study has been conducted in air and in freon-12 in this facility . the wing planform employed had an aspect ratio of 4.0, a taper ratio of 0.6, and 45 of quarter-chord sweepback . a sting-mounted full-span model was tested in addition to three sizes of wall-mounted semispan models . a wide range of mass ratio was covered by the tests in air and by flutter calculations made by the modified strip-analysis method of naca research memorandum l57l10 . a limited amount of data was obtained in freon-12 . results of the tests in air and in freon-12 are in good agreement with the flutter calculations at all mach numbers . the test data compare favorably with previously published transonic flutter data for the same wing planform . the results indicate that flutter characteristics obtained in freon-12 may be interpreted directly as equivalent flutter data in air at the same mass ratio and mach number ." }, { "id": "252", "score": 0.1894, "title": "an investigation of interference effects on similar models of different size in various transonic tunnels in the u.k. .", "abstract": "an investigation of interference effects on similar models of different size in various transonic tunnels in the u.k. . details are given of a programme of tests being made on similar swept-wing models in transonic tunnels of different types . force measurement results at subsonic speeds in the r.a.e. 3 ft. by 3 ft. slotted tunnel show only small interference effects for models of moderate blockage at low incidence., at higher incidences, the interference effect on lift becomes appreciably greater than estimated by theory, and significant pitching moment differences occur, apparently due to wall interference on the wing flow field . comparable but smaller effects are evident in the results from the a.r.a. 9 ft. by 8 ft. perforated tunnel . at speeds just above m = 1, the force fluctuates as speed is increased, because of wave reflection interference . the magnitude of the fluctuations diminishes as speed is further increased and this reduction is more marked in the perforated tunnel . pressure measurements along the top of the body at zero incidence show delay in shock movements at high subsonic speeds indicating a blockage effect on speed., the effect is larger in the perforated tunnel though smaller than predicted by theory . above m = 1, both expansion and shock waves are strongly reflected in the slotted tunnel but considerable alleviation, particularly of shock waves, is achieved in the perforated tunnel, for which an analysis of the effects is given, showing for example, the effect of the open-area distribution of the walls ." }, { "id": "1341", "score": 0.1738, "title": "investigation of wing flutter at transonic speeds for six systematically varied wing plan forms .", "abstract": "investigation of wing flutter at transonic speeds for six systematically varied wing plan forms . an investigation of the effects of systematic variations in wing plan form on the flutter speed at mach numbers between 0.73 and 1.43 has been conducted in the 26-inch langley transonic blowdown tunnel . the angle of sweepback was varied from 0degree to 60degree on wings of aspect ratio 4, and the aspect ratio was varied from 2 to 6 on wings with experimental flutter speed and the reference flutter speed calculated on the basis of incompressible two-dimensional flow . this ratio, designated as the flutter-speed ratio, is plotted as a function of mach number for the various wings . it is found that the flutter-speed ratio increased rapidly past sonic speed for sweep angles of 45degree and less, indicating a favorable effect of mach number . for sweepback of mach number range of the tests . reducing the aspect ratio had a favorable effect on the flutter-speed ratio which was of the order of 100 percent higher for the aspect-ratio-2 wing than for the aspect-ratio-6 wing . this percentage difference was nearly constant throughout the mach number range, indicating that the effect of mach number was about the same for all aspect ratios tested ." }, { "id": "503", "score": 0.1701, "title": "theoretical prediction of the transonic characteristics of airfoils .", "abstract": "theoretical prediction of the transonic characteristics of airfoils . it is shown that the author's transonic-flow airfoil theory can be used to estimate transonic drag-rise and onset-of- separation-effects mach numbers without reference to experimental results . a simple comparative method is applied to a series of airfoils, and the results are analyzed to determine some of the design features of importance in transonic flow . an improvement to this scheme is shown to give results in good agreement with experiment for both the first appearance of shock waves and the onset of separation effects . application to finite swept wings is briefly considered and illustrated ." }, { "id": "810", "score": 0.1616, "title": "the shock wave noise problem of supersonic aircraft in steady flight .", "abstract": "the shock wave noise problem of supersonic aircraft in steady flight . data are presented which provide an insight into the nature of the shock-wave noise problem, the significant variables involved, and the manner in which airplane operation may be affected . flight-test data are also given, and a comparison with the available theory is made . an attempt is also made to correlate the subjective reactions of observers and some associated physical phenomena with the pressure amplitudes during full-scale flight . it is indicated that for the proposed supersonic transport airplanes of the future, booms on the ground will most probably be experienced during the major portion of the flight plan . the boom pressures will be most severe during the climb and descent phases of the flight plan . during the cruise phase of the flight, the boom pressures are of much lesser intensity but are spread laterally for many miles . the manner in which the airplane is operated appears to be significant,. for example, the boom pressures during the climb, cruise, and descent phases can be minimized by operating the airplane at its maximum altitude consistent with its performance capabilities ." }, { "id": "809", "score": 0.1473, "title": "an investigation of the influence of lift on sonic-boom intensity by means of wind tunnel measurements of the pressure fields of several wing-body combinations at a mach number of 2. 01 .", "abstract": "an investigation of the influence of lift on sonic-boom intensity by means of wind tunnel measurements of the pressure fields of several wing-body combinations at a mach number of 2. 01 . an investigation of the effect of lift on sonic-boom intensity has been performed by means of wind-tunnel measurements of the pressure fields surrounding small wing-body combinations . the tests were conducted in the langley 4- by 4-foot supersonic pressure tunnel at a mach number of per foot . effects of lift were found to be real and significant . measured bow-shock intensities agreed fairly well with, but were consistently less than, shock intensities estimated by theoretical methods . available flight data were examined for correlation with wind-tunnel test results ." } ], "milestone": { "id": "arch-0.3-bge", "label": "ARCH-0.3 candidate BGE vector hybrid", "searchEvolutionId": "SE-0003", "architectureVersion": "ARCH-0.3-candidate", "architecture": "bge-vector-hybrid", "status": "candidate-runtime-enabled-with-workers-ai", "resultSummary": { "ndcgAt10": 0.3533, "binaryNdcgAt20": 0.4926 }, "remoteIndex": "cranfield-v0-bge-base-en-v15-gen023", "validatedArtifacts": [ "experiments/cranfield-v0/evaluation-live-opensearch-bge-base-en-v15-gen023.json", "experiments/cranfield-v0/evaluation-live-opensearch-bge-base-en-v15-k20-binary-gen023.json" ] } }