PDF/ocr_dataset_1064.jsonl

{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 24, "total_pages": 47, "image_filename": "19930083221_p24.jpg", "text": "22\nNACA TN No. 1824\n\nenclosing the wing. In the three following sections the latter\napproach will be considered and the limiting value of drag at $M_0=1$\ncomputed. The initial portion of this theory requires the evaluation\nof source and doublet effectiveness at infinity and the concept of\nequivalent source position, an idea which appears to have been given\nfirst by W. D. Hayes in reference 14.\n\nConsider, as in figure 10, a point P with coordinates x, y, z\nlying within the induced field of a supersonic wing. The Mach fore-\ncone from P is given by the relation\n\n$$x-x_1 = \\beta\\sqrt{(y-y_1)^2 + (z-z_1)^2} \\quad (31)$$\n\nwhere $x_1, y_1, z_1$ are running coordinates of a point on the surface\nof the cone. Introducing polar\ncoordinates\n\n$$y = r \\cos \\theta, z = r \\sin \\theta$$\n\nand rewriting the abscissa of\nP in the form\n\n$$x = x_0+\\beta r$$\n\nit follows that the trace of\nthe forecone in the $z_1=0$\nplane is, in the limit as r\napproaches infinity,\n\n$$x_1 = \\beta y_1 \\cos \\theta + x_0 \\quad (32)$$\n\n[Figure: Coordinates used in study of supersonic source.]\n\nIt is, moreover, possible to show that the effect on the velocity\npotential at the point P as r approaches infinity is the same\nfor all points $(x_1, y_1, 0)$ for which $x_1-\\beta y_1 \\cos \\theta = \\text{constant}$. The\nvalue of this effect is\n\n$$\\varphi = \\frac{1}{2\\pi\\sqrt{2\\beta r(x_0-x_1+\\beta y_1 \\cos \\theta)}} \\quad (33)$$\n\nand follows from the asymptotic evaluation of the supersonic source\npotential\n\n$$\\varphi = \\frac{1}{2\\pi\\sqrt{(x-x_1)^2-\\beta^2(y-y_1)^2-\\beta^2 z^2}}$$", "timestamp": "2026-07-22T06:24:06.170563+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 6, "total_pages": 92, "image_filename": "19930085870_p6.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 5\n\nTest methods.- Measurements of lift, drag, and pitching moment were made through an angle-of-attack range of approximately $\\pm 6^\\circ$. With the optical system for indicating angle of attack, the indicated angle may be taken as the true value since the load deflection of the wings ahead of the mirror was found to be negligible. Corrections due to the support deflection have been applied to the moment results in calculation of the moment due to drag.\n\nIn an effort to obtain the order of magnitude of the tare forces on the sting, force measurements were made of the sting alone at the three Mach numbers. The wedge-shaped gap normally occupied by the wing was filled with metal flush with the sting surfaces. Lift and moment of the sting alone were very small, and any effects of the sting on test results are assumed negligible. The drag of the sting alone showed only a very small variation with angle of attack. For the elliptical- or wedge-leading-edge wing having least minimum drag, the drag of the sting alone is approximately 10 percent of the minimum drag. In the wing tests, part of the sting as tested alone is no longer exposed to the air stream, and the remainder of the sting is partially immersed in the boundary layer of the wing. For this reason, the contribution of the sting to the total minimum drag is somewhat less than the 10-percent figure. For the wings having much larger minimum drag, the contribution of the sting may approach values less than 1 percent. With this in mind, the drag results may be compared quantitatively with theory, although no correction for sting drag has been applied.\n\nThere was some doubt as to whether the pressures on either side of the sting within the sting windshield would remain the same if the lips of the windshield were not exactly centered with respect to the sting shoulders. Pressure measurements showed that, provided the lips of the windshield lay behind the sting shoulders, any off-center condition produced no differential in pressure between the sides of the sting and therefore contributed no error to lift-scale measurements. A correction to the drag was applied to account for the difference in pressure between free stream and sting-shield-and-balance enclosing box.\n\nIn the course of the present tests, a liquid-film method for observation of boundary-layer transition, similar to that developed in reference 12 and at the Ames Laboratory (reference 11) was used to supplement the schlieren photographs and pressure distributions. Briefly, the liquid-film method depends upon the greater shear intensity of turbulent boundary layers to vaporize a film of liquid much more rapidly than the comparatively low shear intensity of laminar regions. The ratio of time for drying of the laminar areas to the turbulent areas is approximately 5 to 1 at low Reynolds numbers and greater at high Reynolds numbers; however, it is quite possible for laminar regions very near the leading edge of an airfoil, where the boundary layer is very thin, to show the same drying rates as turbulent areas due to the initial intensity of the shear at the surface. In any case, the shear intensity and the resulting\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:24:07.056494+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 20, "total_pages": 46, "image_filename": "19930085548_p20.jpg", "text": "NACA RM No. E8L30\n19\n\n4. Collins, John H., Jr.: Alterations and Tests of the \"Farnboro\" Engine Indicator. NACA TN No. 348, 1930.\n\n5. Stanitz, John D.: An Analysis of the Factors That Affect the Exhaust Process of a Four-Stroke-Cycle Reciprocating Engine. NACA TN No. 1242, 1947.\n\n6. Schweitzer, P. H.: Porting of Two-Stroke Cycle Diesel Engines - Pt. II. Diesel Power & Diesel Transportation, vol. 20, no. 6, June 1942, pp. 477-479.\n\n7. Taylor, C. Fayette, and Taylor, Edward S.: The Internal Combustion Engine. International Textbook Co. (Scranton, Pa.), 1938, pp. 240-255.\n\n8. Pope, Arthur W., Jr.: KHD Two Cycle Engine Developments with Schnuerle Loop Scavenge System. Fiat Final Rep. No. 683, Off. Director of Intell., U. S. Office Military Govt. (Germany), Jan. 9, 1946. (Available from U. S. Dept Commerce as PB No. 30041.)\n\n9. Pinkel, Benjamin, and Turner, L. Richard: Thermodynamic Data for the Computation of the Performance of Exhaust-Gas Turbines. NACA ARR No. 4E25, 1944.", "timestamp": "2026-07-22T06:24:12.564209+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 25, "total_pages": 62, "image_filename": "19930082918_p25.jpg", "text": "24\nNACA TN 1940\n\nand that the resistance to crack propagation on the other hand is at a maximum. Figures 17 and 20 also show, however, that the plastic strain before fracture decreased with increasing time for rupture. No obvious reason for this presented itself during the investigation covered herein.\n\n### Limitations of Results\n\nThis report is limited to the presentation of a method of determining the fundamental mechanisms by which processing, heat treatment, and chemical composition control the properties of alloys at high temperatures. A relatively limited amount of data for solution-treated and aged low-carbon N-155 alloy has been obtained and interpreted in terms of the proposed method. The resulting theories require extension and improvement from similar investigations on many alloys, as well as from more test conditions on low-carbon N-155 alloy. It is believed, however, that the approach to the problem is reasonably sound and that the limiting factors to a general theory of the metallurgical factors controlling high-temperature strength of alloys are the large volume of testing required and development of suitable experimental techniques.\n\nIn regard to low-carbon N-155 alloy, there are obviously numerous important aspects of the problem which have not been adequately covered by this report. Of primary importance is the fact that the structural measurements have not been correlated with long-time creep and rupture strengths at 1200° F or with any time period at other temperatures. Certain refinements in X-ray techniques seem desirable. Information is also needed regarding the effect of time and deformation during testing to assess the reliability of the assumption that such changes had little effect on known initial structure. And, finally, it would be especially desirable to know the exact composition of the precipitating phases. The amount of time required for development of suitable techniques and the volume of experimental work have been the limiting factors to date.\n\n### CONCLUSIONS\n\nAn experimental procedure is described which is believed suitable for establishing the fundamental mechanisms by which processing, heat treatment, and chemical composition control the properties of alloys at high temperatures. This method relates microstructures and X-ray diffraction characteristics; after various prior treatments, to creep and rupture test properties.\n\nApplication of this method to solution-treated and aged low-carbon N-155 alloy and correlation with the short-time creep and rupture.", "timestamp": "2026-07-22T06:24:13.166750+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 35, "total_pages": 44, "image_filename": "19930082566_p35.jpg", "text": "```markdown\nNACA TN NO. 1889\n\nMaximum principal stress, $\\sigma_1'$, psi\n\nN, cycles\n\n(c) For stress ratio $R = \\sigma_2' / \\sigma_1' = 1$.\n\nFigure 12.- Continued.\n\n33\n\n[Figure: A graph plotting Maximum principal stress, $\\sigma_1'$, psi (y-axis, $0$ to $50 \\times 10^3$) against N, cycles (x-axis, $4 \\times 10^4$ to $10^7$). The graph shows a curve starting at approximately $(10^5, 23 \\times 10^3)$ and decreasing to a plateau around $12 \\times 10^3$ psi for $N > 10^6$. Data points are marked with circles. Two arrows point to the right from the plateau region. The NACA logo is in the bottom right corner of the plot area.]\n```", "timestamp": "2026-07-22T06:24:16.850613+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 19, "total_pages": 31, "image_filename": "19930085859_p19.jpg", "text": "```markdown\nNACA RM No. L9B25\n\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nM\n1.18 $\\circ$\n1.15 $\\diamond$\n1.10 $\\square$\n1.08 $\\nabla$\n1.05 $\\nabla$\n1.03 $\\nabla$\n1.00 $\\circ$\n.98 $\\diamond$\n.95 $\\diamond$\n.93 $\\square$\n.90 $\\triangle$\n.88 $\\triangle$\n.85 $\\triangle$\n.80 $\\diamond$\n.70 $\\square$\n.60 $\\circ$\n\nPitching-moment coefficient, $C_m$\nLift coefficient, $C_L$\nM\n1.18 $\\circ$\n1.15 $\\diamond$\n1.10 $\\square$\n1.08 $\\nabla$\n1.05 $\\nabla$\n1.03 $\\nabla$\n1.00 $\\circ$\n.98 $\\diamond$\n.95 $\\diamond$\n.93 $\\square$\n.90 $\\triangle$\n.88 $\\triangle$\n.85 $\\triangle$\n.80 $\\diamond$\n.70 $\\square$\n.60 $\\circ$\n\nDrag coefficient, $C_D$\nLift coefficient, $C_L$\n[Figure: NACA logo]\n\nFigure 7.- Wing-alone aerodynamic characteristics for a model with $35^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n17\n```", "timestamp": "2026-07-22T06:24:22.553195+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 40, "total_pages": 50, "image_filename": "19930082592_p40.jpg", "text": "NACA TN 1914\n39\n\nOxidation front\nGrain-boundary\noxide penetration\n\n<!-- Image (76, 169, 897, 735) -->\n\nNACA\nC-22913\n2-7-49\n\nFigure 14. - Oxidation interface of 30-percent-tungsten - titanium carbide ceramal.\nOxidation proceeds along grain boundaries rather than moving as linear front. Temperature, 1785° F; time at temperature, 30 hours; unetched; magnification, X750.", "timestamp": "2026-07-22T06:24:29.086856+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 1, "total_pages": 29, "image_filename": "19930085899_p1.jpg", "text": "RM No. L9A21\n\nNACA RESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS OF A WING WITH QUARTER-CHORD LINE SWEPT BACK $45^\\circ$, ASPECT RATIO 4, TAPER RATIO 0.6, AND NACA 65A006 AIRFOIL SECTION\n\nTRANSONIC-BUMP METHOD\n\nBy\n\nJoseph Weil and Kenneth W. Goodson\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nFebruary 24, 1949\nDeclassified September 27, 1954", "timestamp": "2026-07-22T06:24:29.377014+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 31, "total_pages": 46, "image_filename": "19930082613_p31.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:24:30.890178+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 37, "total_pages": 65, "image_filename": "19930082546_p37.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:24:31.144719+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 4, "total_pages": 26, "image_filename": "19930085890_p4.jpg", "text": "NACA RM No. E9C11\n\nwere designed to produce solid streams of propellants. In injection plate A, which was used for most of the experiments, four diborane injectors and four liquid-oxygen injectors were arranged in intersecting pairs. The four diborane injectors were on an inner circle and directed diborane in streams parallel to the engine axis. The four liquid-oxygen injectors were on an outer circle and directed oxygen in streams to intersect the respective diborane streams. Injection plate B had another arrangement of injectors. Two diborane injectors and two oxygen injectors directed streams that intersected at a common point on the axis of the engine. Each injection plate had a pressure tap to obtain combustion pressure.\n\nTwo combustion chambers were used to obtain two values of characteristic length $L^*$ (ratio of combustion-chamber volume to exhaust-nozzle-throat area) and differed essentially only in length. Both combustion chambers had an inner diameter of $2\\frac{1}{4}$ inches; chamber A had a total length of $18\\frac{1}{2}$ inches and an $L^*$ of 325 inches; chamber B had a length of $6\\frac{3}{8}$ inches and $L^*$ of 159 inches. Combustion chamber A was used for most of the experiments. Both chambers were machined from forged electrolytic copper, had a wall thickness of 5/8 inch, and were chrome-plated on the inner surface. Each combustion chamber was equipped with a pressure tap with a thick-wall copper connection.\n\nThe convergent-divergent nozzle was machined from forged electrolytic copper and was chrome-plated on the inside. The throat diameter was 0.549 inch, the total divergence angle was $30^\\circ$, and the exit diameter was 1.043 inches. The nozzle was designed to obtain optimum performance for an expansion pressure ratio of 20.4 with a ratio of specific heats of the exhaust gases of 1.2.\n\nInstrumentation\n\nPropellant flow. - The oxidant and the fuel tanks were each suspended from the lever arms of counterbalances. The balances used were standard beam balances each modified to permit suspension of a propellant tank and a counterbalance tank. The counterbalance tank was a container of water that could be remotely filled or emptied as required to balance initially the propellant and the propellant tank. Dashpots were installed on the counterbalance end of the beams to dampen oscillations introduced during the runs. Unbalanced forces caused by change of weight in the", "timestamp": "2026-07-22T06:24:35.847120+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 57, "total_pages": 98, "image_filename": "19930086073_p57.jpg", "text": "NACA RM A59I04\n\nLift coefficient, $C_L$\n\n$\\beta$, deg\n○ 0.0\n□ 6.0\n◇ 12.0\n△ 15.9\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 11.— Concluded.\n\n55", "timestamp": "2026-07-22T06:24:36.106424+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 22, "total_pages": 149, "image_filename": "19930083192_p22.jpg", "text": "```markdown\n18\nNACA TN 1976\n\nAs indicated in table II, a variety of weather conditions, from line squalls to gustiness in the ground boundary layer, is included in the present data. If the data of figure 12(a) are assumed to be comparable, regardless of airplane, the type of weather appears to have no effect on the gust-gradient distance.\n\nFigure 14 indicates that the average gust-gradient distance is independent of altitude in convective clouds. The data scatter somewhat but the over-all agreement appears to be good.\n\nGeneral considerations would indicate that the gradient distance might increase with gust velocity. Previous study and figure 12(a) have indicated such a relation, although the evidence is not conclusive. When the faired curves of figures 12(a) and 16 are considered, such a variation appears to be a reasonable estimate. The evidence available (fig. 12(a)) is believed to indicate that for large gusts the average gradient distance is essentially constant but decreases rapidly at the smaller gust velocities.\n\nSpanwise gust distribution.- The shapes of the lateral gust distributions, as indicated by figure 15, take many irregular forms. From consideration of the frequency of occurrence, the triangular, double-triangle, and unsymmetrical gust shapes predominate. Inasmuch as the double triangle can lead to wing bending moments along the span equal to those due to a uniform gust distribution, while the triangular gust shape would lead to reduced wing bending moments, the selection for design of a uniform gust velocity across the span for the symmetrical load condition is conservative. The relative percentages of the symmetrical (triangle, and so forth) and unsymmetrical gust shapes indicate that the unsymmetrical gust varying uniformly across the span appears quite frequently.\n\nComparison of the data for the average lateral gust-gradient distance as a function of gust velocity U (fig. 16) with corresponding results for longitudinal gust-gradient distances (fig. 12(a)) indicates that the gradient distances are essentially the same in both directions. This agreement is indicated by the fact that the same empirical curve fits both sets of data and the results show the same trend of increasing gust-gradient distance with increasing gust velocity. Because data from only one airplane have been presented and the precision of measurement of lateral values is poor, as previously cited, the excellent agreement between figures 16 and 12(a) is to some degree fictitious, the degree being unknown. Within the limitations of results the lateral and longitudinal gust dimensions are concluded to be essentially the same.\n\nInspection of figure 18, relating the unsymmetrical components of the effective gust velocity to the average uniform gust velocity across the span, indicates that no clear relations between the unsymmetrical\n```", "timestamp": "2026-07-22T06:24:38.059012+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 13, "total_pages": 60, "image_filename": "19930085862_p13.jpg", "text": "NACA RM No. L9A07\n\nthat the variation of hinge-moment coefficient with aileron deflection becomes more negative at large deflections, it should be noted that the aileron would be more underbalanced at these large deflections than is indicated in figure 16.\n\nNormal-force characteristics.- The aileron normal-force coefficients presented in figures 5 to 10 represent the forces on the aileron behind the hinge line. In the use of these data in the design of an aileron with a sealed internal balance, account must be taken of the additional forces acting on the balance. The maximum values of the aileron normal-force coefficients were about the same for the flapped or unflapped wing configurations. The stall studies presented in figure 17 show that the aileron on the plain wing is completely stalled at an angle of attack of $16.9^\\circ$, whereas the aileron on the flapped wing is only partly stalled at angles of attack of more than $19^\\circ$. As a result of this early aileron stall the variation of aileron normal-force coefficient with angle of attack for the plain wing was not as linear as that for the wing with the leading-edge devices.\n\nSpoiler Characteristics\n\nRepresentative data obtained from tests of numerous wing and spoiler configurations are presented in figures 18 to 21.\n\nRolling-moment characteristics.- It is apparent from the data presented in figures 18 to 21 that the origin and progression of the stall are reflected in the rolling-moment coefficients contributed by the various spoiler arrangements. In the case of the plain wing (fig. 18(a)), the outboard section of the $0.775\\frac{b}{2}$ spoiler is enveloped in tip stall at a relatively low angle of attack (approximately $8.6^\\circ$; fig. 17) which results in an abrupt decrease in $C_l$. The same abrupt decrease in $C_l$ is indicated for a $0.375\\frac{b}{2}$ spoiler located at the tip, whereas a $0.4\\frac{b}{2}$ spoiler located inboard of the $0.60\\frac{b}{2}$ station does not encounter the effects of the wing stall until an angle of attack of approximately $12^\\circ$. The wing equipped with the high-lift and stall-control devices exhibited an initial stalled region behind the inboard ends of the leading-edge flaps, and as the spoilers extended into this region there was a marked reduction in $C_l$ (figs. 20 and 21(a)). It should be noted that for the flapped configurations some rolling-moment coefficient is produced for angles of attack corresponding to $C_{L_{max}}$.", "timestamp": "2026-07-22T06:24:40.852635+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 7, "total_pages": 92, "image_filename": "19930085870_p7.jpg", "text": "6\nCONFIDENTIAL\nNACA RM No. L9D07\n\nrate of energy dissipation in the particular region will determine whether the region remains wet or dry and conclusions reached from liquid-film methods are made on this basis. The models were given a matte black finish before applying the liquid-film solution. Upon completion of a run, the models were dusted with powder. Accordingly, the wet regions appear white in the photographs and the dry regions remain black.\n\nAll schlieren photographs were taken with the knife-edge horizontal. At the time the tests of the elliptical leading-edge series were conducted, the spark system normally used for the schlieren apparatus was inoperative and a manual shutter was substituted. This explains the poor resolution of unsteady flows evident on the schlieren photographs of these wings for which the exposure time of 1/100 second was quite large in comparison with the several microseconds for the spark exposures.\n\nPrecision of data.- The estimated probable errors in the aerodynamic quantities are included in the following table: The value of $\\pm 0.08^\\circ$ given for angle of attack is a result of error in the initial referencing of each wing with respect to stream direction. The value of $\\pm 0.01^\\circ$ is the error that might be incurred in relative-angle-of-attack readings for a given test.\n\n| M | $C_L$ | $C_D$ | $C_m$ | M | $\\alpha$ (deg) | | R |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | Initial | Relative | |\n| 1.62 | $\\pm 0.0004$ | $\\pm 0.0004$ | $\\pm 0.0018$ | $\\pm 0.01$ | $\\pm 0.08$ | $\\pm 0.01$ | $\\pm 20,000$ |\n| 1.92 | | | | | | | |\n| 2.40 | | | | | | | |\n\nReynolds numbers of tests.- The test values of the Reynolds numbers based on $\\bar{c}$, two-thirds of the root chord, are given in the following table:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:24:58.282010+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 25, "total_pages": 37, "image_filename": "19930082646_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:25:07.778913+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 26, "total_pages": 26, "image_filename": "19930085485_p26.jpg", "text": "```markdown\n24\nNACA RM No. L8K02\n\nCONFIDENTIAL\n\nRolling-moment coefficient, $C_{l\\delta}$\n\nAileron deflection, $\\delta_a$, deg\n\n| | $M = .5$ | $M = .8$ |\n| :--- | :---: | :---: |\n| | [Graph] | [Graph] |\n| | | |\n| | $M = 1.0$ | $M = 1.1$ |\n| | [Graph] | [Graph] |\n\nAileron profile(t)\n\n| | | |\n| :--- | :--- | :--- |\n| — | .0 | Cir. arc |\n| — | .37 | Flat-sided |\n| — | .50 | |\n| - - - | -1.00 | |\n\nNACA\n\nFigure 19.- The effect of aileron profile on the rolling-moment characteristics. $\\alpha = 0^\\circ$; data from transonic bump.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:25:20.414901+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 21, "total_pages": 46, "image_filename": "19930085548_p21.jpg", "text": "Reduction\ngear\n\nTurbine\n\nTwo-stroke-cycle\ncompression-ignition\nengine\n\nGear\n\nCompressor\n\n[Figure: Diagrammatic sketch of gas-generator engine used in analysis (reference 1).]\n\nFigure 1. - Diagrammatic sketch of gas-generator engine used in analysis (reference 1).\n\nNACA\n\nNACA RM No. E8L30\n\n20\n\n1077", "timestamp": "2026-07-22T06:25:22.358565+00:00"}
{"citation_id": "19930085859", "source_url": "https://ntrs.nasa.gov/api/citations/19930085859/downloads/19930085859.pdf", "page_number": 20, "total_pages": 31, "image_filename": "19930085859_p20.jpg", "text": "18\nNACA RM No. L9B25\n\nBending-moment coefficient, $C_B$\nLift coefficient, $C_L$\n\n| M | |\n| :--- | :--- |\n| 1.18 | $\\circ$ |\n| 1.15 | $\\diamond$ |\n| 1.10 | $\\square$ |\n| 1.08 | $\\triangledown$ |\n| 1.05 | $\\nabla$ |\n| 1.03 | $\\nabla$ |\n| 1.00 | $\\triangle$ |\n| .98 | $\\diamond$ |\n| .95 | $\\diamond$ |\n| .93 | $\\square$ |\n| .90 | $\\square$ |\n| .88 | $\\triangle$ |\n| .85 | $\\triangle$ |\n| .80 | $\\diamond$ |\n| .70 | $\\square$ |\n| .60 | $\\circ$ |\n\n[Figure: A graph plotting Bending-moment coefficient ($C_B$) against Lift coefficient ($C_L$). The vertical axis ranges from -0.16 to 0.32. The horizontal axis ranges from -0.2 to 0.8. The graph contains multiple series of data points represented by different symbols (circles, diamonds, squares, triangles, inverted triangles) corresponding to different Mach numbers (M) listed in the legend on the right. A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 7.- Concluded.", "timestamp": "2026-07-22T06:25:24.403319+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 41, "total_pages": 50, "image_filename": "19930082592_p41.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:25:25.432054+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 23, "total_pages": 149, "image_filename": "19930083192_p23.jpg", "text": "```markdown\nNACA TN 1976\n19\n\ncomponents and the average gust velocity exists but that two boundaries\nappear to exist. A fairly definite boundary rising from the lower left-\nhand corner of the figure and corresponding to equality of the unsym-\nmetrical component of gust velocity and the average gust velocity is\nindicated. As higher values of the average gust velocity are considered,\nthe results scatter considerably, but an upper boundary is found at\nfrom 9 to 10 feet per second. The boundary holds for gust velocities\nup to 40 to 48 feet per second. On the basis of this upper limit, the\nunsymmetrical component of gust velocity above 10 feet per second can be\nconsidered rare. Therefore, the unsymmetrical gust can be considered to\nbe composed of two components, an unsymmetrical component amounting to\na maximum of about 10 feet per second with opposite signs imposed at each\nwing tip and a linear gradient across the span, and an average uniform\ngust velocity of any pertinent value.\n\nA question of some importance is the determination of the magnitude\nof the unsymmetrical components for other aircraft of entirely different\nspans. For example, if the span of the airplane were doubled, would the\nunsymmetrical component amount to twice the value obtained from the data\nshown herein? In order to obtain information on this question, the\nmaximum value of the unsymmetrical gust components computed in refer-\nence 13 from angular-acceleration data was compared with the values\nobtained herein. It was found that for the XB-15 airplane, which has a\nspan of about 150 feet, the unsymmetrical components of gust velocity\nat the wing tip would be about 13.7 feet per second. This value is\nslightly greater than indicated by figure 18 but is not, by any means,\nproportional to the span of the airplane. Similar data of a statistical\nnature obtained for the XC-35 airplane (table V) indicate that the maxi-\nmum value of the angular acceleration on that airplane varied from about\n1.95 to about 2.25 radians per second per second. Computations utilizing\nthe formula given in reference 13 yield a maximum gust-velocity component\nof about $12\\frac{1}{2}$ feet per second. This value is in fair agreement with the\none obtained for the XB-15 airplane and indicates that the tip values\nare independent of span. Since the values based on angular accelera-\ntion are somewhat more conservative than those based on wing pressures,\nmore reliance should be placed on these values.\n\nWhen the information on unsymmetrical gusts is utilized, the average\neffective gust velocity that should be used in conjunction with the maxi-\nmum unsymmetrical component of 14 feet per second must be determined.\nThe data of figure 19 indicate that for an effective gust velocity in the\nneighborhood of 30 feet per second, the average gust velocity for the\nunsymmetrical gust would be about 22 feet per second.\n\nGust spacing.- Inspection of table IV indicates that the spacing of\nrepeated gusts for two or three gusts in sequence is about 22 chords\nwith a spread in actual values of some 20 chords for all sequences.\n```", "timestamp": "2026-07-22T06:25:26.859428+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 25, "total_pages": 47, "image_filename": "19930083221_p25.jpg", "text": "NACA TN No. 1824\n23\n\nfor large values of r. The potential at P for the source at\n$(x_1, y_1, 0)$ is thus the same as for the source shifted along the\ntrace to $(x_1-\\beta y_1 \\cos \\theta, 0, 0)$, the intercept of the trace on the x\naxis. For Mach numbers near one, equation (33) can be rewritten\n\n$$\n\\varphi = \\frac{1}{2\\pi \\sqrt{2\\beta r(x_0-x_1)}} \\quad (34)\n$$\n\nand is equivalent to the potential at P for the source at $(x_1, 0, 0)$.\nThe induced velocities at P due to a source at $(x_1, y_1, 0)$ follow\nimmediately, for arbitrary $M_0$ and for $M_0$ near one, from the\ngradients of $\\varphi$ in equations (33) and (34). It is important to note\nthat equation (33) is a function of the azimuthal angle of P so\nthat, in general, a source does not have a fixed equivalent position\nwith respect to its potential at infinity; equation (34), however,\nis independent of the azimuth $\\theta$.\n\nThe source-sink potential is applicable to the study of\nsymmetrical nonlifting wings. When lifting surfaces are to be analyzed,\nthe doublet potential\n\n$$\n\\varphi = \\frac{\\beta^2 z}{2\\pi [(x-x_1)^2 - \\beta^2(y-y_1)^2 - \\beta^2 z^2]^{3/2}}\n$$\n\nmust be considered and the question of equivalent doublet position\nwith respect to the potential at infinity arises. In this case the\ndoublet position can again be shifted parallel to the trace of the\nMach cone from P at infinity and the potential at P is given by\nthe expression\n\n$$\n\\varphi = \\frac{\\beta^2 z}{2\\pi [2\\beta r(x_0-x_1+\\beta y_1 \\cos \\theta)]^{3/2}} \\quad (35)\n$$\n\nand, for Mach numbers near one,\n\n$$\n\\varphi = \\frac{\\beta^2 z}{2\\pi [2\\beta r(x_0-x_1)]^{3/2}} \\quad (36)\n$$\n\nMomentum relations.— The vectorial force $\\vec{F}$ on an aerodynamic\nbody inside a control surface S is given by the surface integral", "timestamp": "2026-07-22T06:25:29.330141+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 32, "total_pages": 46, "image_filename": "19930082613_p32.jpg", "text": "NACA TN 1938\n31\n\n[Figure: Cross-sectional view of a material sample with a circled area labeled \"Detail A\"]\n\n(a) Cross-sectional view; magnification, X100.\n\n[Figure: Detailed view of \"Detail A\" showing various features with labels: \"Tubular branch\", \"Subsurface scale\", \"Metal\", \"Surface scale\", \"Subsurface scale\", and \"Main crack\". A NACA stamp is visible in the bottom left corner with the text \"NACA C-22629 12-9-48\".]\n\n(b) Detail A showing surface scale at edge of crack; magnification, X1500.\n\nFigure 9. - Crack extending from punched edge. Type-A liner; etchant, none; condition, failed in service.", "timestamp": "2026-07-22T06:25:33.665210+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 2, "total_pages": 29, "image_filename": "19930085899_p2.jpg", "text": "NACA RM No. L9A21\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS OF A WING WITH QUARTER-CHORD\nLINE SWEPT BACK $45^\\circ$, ASPECT RATIO 4, TAPER RATIO 0.6,\nAND NACA 65A006 AIRFOIL SECTION\n\nTRANSONIC-BUMP METHOD\n\nBy Joseph Weil and Kenneth W. Goodson\n\nSUMMARY\n\nAs part of an NACA transonic research program, a series of wing-body\ncombinations are being investigated in the Langley high-speed 7- by 10-\nfoot tunnel up to a Mach number of 1.18 utilizing the transonic bump.\n\nThis paper presents the results of the investigation of a wing-\nfuselage combination employing a wing with the quarter-chord line swept\nback $45^\\circ$, with aspect ratio 4, taper ratio 0.6, and an NACA 65A006 air-\nfoil section. Lift, drag, pitching moment, and root bending moment were\nobtained for the wing-alone and wing-body configurations. Effective down-\nwash angles and dynamic-pressure characteristics in the region of the tail\nplane were also obtained and are presented for a range of tail heights at\none tail length. The effects of two wing-fence arrangements were investi-\ngated. In order to expedite publishing of these data only a brief analysis\nis included.\n\nINTRODUCTION\n\nThe urgent need for aerodynamic design data in the transonic speed\nrange has led to the establishment of a special NACA committee for tran-\nsonic research. As part of the NACA transonic research program recom-\nmended by this committee a series of wing-body configurations having wing\nplan form as the chief variable are being investigated in the Langley high-\nspeed 7- by 10-foot tunnel utilizing the transonic-bump test technique.\nFor each wing-fuselage combination investigated the lift, drag, pitching-\nmoment, and root bending-moment characteristics are determined up to a\nMach number of about 1.18. In addition, effective downwash angles and\ndynamic-pressure characteristics are obtained for a range of tail heights\nat one tail length.", "timestamp": "2026-07-22T06:25:37.099307+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 26, "total_pages": 62, "image_filename": "19930082918_p26.jpg", "text": "NACA TN 1940\n25\n\ncharacteristics at 1200° F indicated the following possible fundamental explanations for the effect of aging on the 1200° F properties.\n\n1. Aging of solution-treated low-carbon N-155 resulted in progressive lowering of the initial (short-time) creep resistance through the removal from solid solution of large-radius or interstitial atoms by precipitation. No optimum precipitate dispersion or state occurred for optimum creep resistance; rather the alloy can be considered as obtaining its optimum creep strength through \"modulation\" of the lattice with the large or small precipitate atoms when they are in the random, or at most nucleated, distribution of the solution-treated state.\n\n2. Short-time aging of solution-treated low-carbon N-155 apparently resulted in a marked increase in short-time rupture strengths over that for unaged material through the growth of a grain boundary phase, this phase acting to strengthen the boundary areas to eliminate intergranular cracking and consequent low resistance to crack propagation. Long-time aging resulted in little further change in short-time rupture strength, and longer-time service at 1200° F was sufficient to develop a grain boundary phase in the unaged material and thus to raise its rupture strength to compare favorably with the strengths of prior-aged material.\n\n3. Because the effect of aging in general was to lower the creep resistance and to raise the rupture strength, for the time periods considered, aging resulted in a material which exhibited greater ductility before fracture.\n\nIt is indicated that alloys might be developed with strength comparable to that of N-155 by the use of other alloying elements, or that an alloy of the same general type with improved creep strength might be developed by replacing or supplementing the elements of large or small atomic radius present in low-carbon N-155. Elements of the same atomic radius as molybdenum, columbium, or tungsten, which could possibly act as substitutes or supplements, include aluminum, silver, and tantalum and elements of larger atomic radius suitable for additional alloying include zirconium, cerium, and titanium. Boron appears to be the only promising alloying element of small atomic radius not at present used in N-155.\n\nIt is emphasized that the foregoing explanations for the effect of aging on the properties of low-carbon N-155 alloy at 1200° F apply at present only to that alloy and are not to be taken as general. It is entirely possible that this alloy will prove to be an unusual one exhibiting behavior which is the exception to some general rule to be established as a result of further investigation on other alloys.\n\nUniversity of Michigan\nAnn Arbor, Mich., January 18, 1949", "timestamp": "2026-07-22T06:25:37.281934+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 14, "total_pages": 60, "image_filename": "19930085862_p14.jpg", "text": "12\nNACA RM No. L9A07\n\nThe effects of spoiler projection or height for the plain wing may be seen by a comparison of figures 18(a) and 19(a). In the low angle-of-attack range, the 0.10c spoiler is several times as effective as the 0.05c spoiler which in all probability is due to the fact that a smaller percentage of the 0.10c spoiler is in the boundary-layer air. At the angles of attack where the boundary layer becomes thicker and flow separation occurs, the effectiveness of both the 0.10c and 0.05c spoilers becomes equal until finally both have zero effectiveness as all the spoiler segments are enveloped in the stalled region.\n\nIn figures 22 to 24 a summary is presented of all spoiler combinations tested. It can be seen in figure 22 that for a given spoiler span on the plain wing the inboard location provided slightly greater values of $C_l$ at low angles of attack than did the outboard location. It is quite possible that the inboard spoilers on a sweptback wing can, due to crossflow, cause spoiling of the flow over sections of the wing outboard of the spoilers. For the flaps-deflected configurations, a spoiler located on the outboard portion of the wing produced higher values of $C_l$ than a spoiler of equal span located inboard. A spoiler of the same span but with its inboard end located at the wing root might result in yet different results. It seems, therefore, that the optimum spanwise spoiler location on a sweptback wing is largely dependent upon the span loading and/or the spanwise center of pressure of that particular wing.\n\nIt can be seen in figure 22 that for the plain wing equipped with a short span of the 0.05c spoilers some rolling-moment reversal was encountered. It is possible that with a short span of the 0.10c spoilers reversal might also be encountered. For this reason no attempt has been made to fair the curves of figures 22 to 24 through $\\frac{v_a}{b/2} = 0$.\n\nOther aerodynamic characteristics.- As indicated in figures 18 to 21, the yawing-moment coefficients obtained with the spoilers on the plain wing were favorable up to an angle of attack of about 12°. Above this angle, adverse yawing-moment coefficients were obtained although the values were small. The addition of leading-edge and trailing-edge flaps resulted in favorable yawing-moment characteristics up to almost the angle of attack for $C_{L_{max}}$. The values of $C_n$ at the lower angles of attack were somewhat larger for the wing with flaps than for the plain wing.\n\nThe maximum changes in the pitching-moment coefficient resulting from the 0.10c projection of the spoilers were 50 to 100 percent greater than those resulting from the maximum deflection of a set of ailerons. The outboard spoilers on the plain wing configuration caused a large positive shift in the pitching-moment curve up to the angle of attack at which tip stalling began. As the separated flow at higher angles of", "timestamp": "2026-07-22T06:25:37.592611+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 38, "total_pages": 65, "image_filename": "19930082546_p38.jpg", "text": "NACA TN No. 1870\n37\n\n[Figure: A black and white photograph showing a man standing next to a large propeller and a wooden structure. A label on the structure reads \"NACA L-55800\".]\n\n(a) Reinforced plywood wall (front view).\nFigure 5.- Simulated fuselage walls used in tests.", "timestamp": "2026-07-22T06:25:43.435867+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 51, "total_pages": 114, "image_filename": "19930086061_p51.jpg", "text": "NACA RM L9J07\n47\n\nLeft semispan\nRight semispan\nUpper\nLower\n(c) $\\psi = 20^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n(d) $\\psi = 35^\\circ$\n\nFigure 12.- Concluded.", "timestamp": "2026-07-22T06:25:46.131086+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 36, "total_pages": 44, "image_filename": "19930082566_p36.jpg", "text": "```markdown\n34\n\n50 x 10³\nMaximum principal stress, $\\sigma_1'$, psi\n40\n30\n20\n10\n0\n4 x 10⁴ 10⁵ N, cycles 10⁶ 10⁷\n[Figure: Graph showing Maximum principal stress vs. N, cycles with data points and a curve]\nNACA\n(d) For stress ratio R = $\\sigma_2' / \\sigma_1'$ = 0.5.\nFigure 12.- Concluded.\nNACA TN NO. 1889\n```", "timestamp": "2026-07-22T06:25:51.345809+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 1, "total_pages": 33, "image_filename": "19930085900_p1.jpg", "text": "NACA PM 19D20\nFILE COPY\nNO 6\nCONFIDENTIAL\nCopy 226\nRM L9D20\n\nNACA\nRESEARCH MEMORANDUM\n\nTHE EFFECT OF AIR-JET AND STRIP MODIFICATIONS ON THE\nHYDRODYNAMIC CHARACTERISTICS OF THE STREAMLINE\nFUSELAGE OF A TRANSONIC AIRPLANE\n\nBy\nBernard Weinflash, Kenneth W. Christopher,\nand Charles L. Shuford, Jr.\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nWASHINGTON\nJune 3, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:25:54.338760+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 22, "total_pages": 46, "image_filename": "19930085548_p22.jpg", "text": "NACA RM No. E8L30\n21\n\n1077\n\nFuel injector\n\nInlet\n\nExhaust\n\nInlet\n\nExhaust\n\nCross section through ports\n\nNACA\n\nFigure 2. - Cross-sectional views of experimental engine.\n\n154-387-A", "timestamp": "2026-07-22T06:25:54.544097+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 42, "total_pages": 50, "image_filename": "19930082592_p42.jpg", "text": "NACA TN 1914\n41\n\n[Figure: Micrograph showing grain-boundary oxide penetration and unoxidized grain boundary]\n\nFigure 15. - Grain-boundary oxide penetration into 30-percent-tungsten - titanium carbide ceramal. Temperature, $1785^\\circ$ F; time at temperature, 30 hours; etchant, potassium hydroxide plus potassium ferricyanide KOH+K$_3$Fe(CN)$_6$; magnification, X750.", "timestamp": "2026-07-22T06:26:00.078846+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 5, "total_pages": 26, "image_filename": "19930085890_p5.jpg", "text": "```markdown\n4\nNACA RM No. E9C11\n\npropellant systems during a run were transmitted to cantilever\nbeams equipped with strain gages. The cantilever beam was designed\nfor a change in deflection force of 4 pounds.\n\nThe strain gages were connected in a resistance-bridge circuit\nof a continuous-recording self-balancing potentiometer. In order\nto permit full-scale deflection of the recorder for different quan-\ntities of propellants used, resistances could be added or removed\nfrom the strain-gage circuit by switches. The potentiometer had an\naccuracy of 0.5 percent of full scale. Weight calibrations made\nbefore and after each run had a maximum variation of 2 percent.\n\nThrust. - The thrust produced by the engine was measured by\nmeans of a cantilever beam equipped with strain gages. The strain\ngages were connected in a resistance-bridge circuit and the unbal-\nance in the circuit, corresponding to the thrust, was recorded on\na continuous-recording, self-balancing potentiometer with an accu-\nracy of 0.2 percent of full scale. Weight calibrations of thrust\nwere made periodically and had a maximum variation of 0.5 percent\nof the total weight.\n\nPressure. - Combustion pressure was measured by Bourdon-tube-\ntype pressure recorders having an accuracy of better than 1 percent\nof full scale. During the experiments, some difficulties were\nexperienced with pressure-tap burnouts and clogging of the tap by\nthe solid products of combustion. In order to insure a combustion-\npressure record for every operation, two measurements of combustion\npressure were made, one from a water-cooled tap in the injection\nplate and the other from a tap with a thick-wall copper connection\nin the combustion chamber. For the second measurement, a small\nhelium bleed was provided to prevent clogging of the tap by solid\ncombustion products. Other pressure measurements made were\npropellant-tank and helium-supply pressure.\n\nTemperature. - Chromel-alumel thermocouples located in the\nwalls of the various exhaust nozzles and combustion chambers used\nin the runs measured the temperature of the metal. The tempera-\ntures were recorded with an accuracy of more than 1.0 percent of\ntheir value on continuous self-balancing, strip-chart potentiometers.\n\nPropellants. - The propellants used in the experiments were\ncommercial liquid oxygen and liquid diborane, which contained\n5 percent impurities, the impurities probably being ethane and\nethyl ether.\n```", "timestamp": "2026-07-22T06:26:03.654662+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 26, "total_pages": 37, "image_filename": "19930082646_p26.jpg", "text": "NACA TN 1980\n25\n\n[Figure: Four rows of two photographs each showing a seaplane on water during take-off. The left column is labeled (a) and the right column is labeled (b).]\n\n$\\tau = 3.1^\\circ$\nV = 25.9 mph\n$\\tau = 6.1^\\circ$\n\n$\\tau = 3.2^\\circ$\nV = 28.0 mph\n$\\tau = 6.0^\\circ$\n\n$\\tau = 3.2^\\circ$\nV = 30.2 mph\n$\\tau = 6.4^\\circ$\n\n$\\tau = 3.3^\\circ$\nV = 32.3 mph\n$\\tau = 6.7^\\circ$\nNACA\nL-59838\n\n(a) Warped forebody and\nextended afterbody.\n\n(b) Basic forebody and\nbasic afterbody.\n\nFigure 12.- Spray in propellers during take-off at design gross load.\n$\\delta_e = -10^\\circ$.", "timestamp": "2026-07-22T06:26:09.151361+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 39, "total_pages": 65, "image_filename": "19930082546_p39.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:26:10.216248+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 8, "total_pages": 92, "image_filename": "19930085870_p8.jpg", "text": "NACA RM No. L9D07 CONFIDENTIAL 7\n\n| Wing | Reynolds number | | |\n| :--- | :--- | :--- | :--- |\n| | **M = 1.62** | **M = 1.92** | **M = 2.40** |\n| 1 | $1.39 \\times 10^6$ | $1.25 \\times 10^6$ | $1.00 \\times 10^6$ |\n| 2 | 1.39 | 1.25 | 1.00 |\n| 3 | 1.38 | 1.23 | .99 |\n| 4 | 1.20 | 1.08 | .86 |\n| 5 | 1.08 | .96 | .77 |\n| 6 | 1.00 | .90 | .72 |\n| 7 | .94 | .84 | .67 |\n| 8 | .86 | .77 | .62 |\n| 9 | .78 | .70 | .56 |\n| 10 | .74 | .66 | .53 |\n| 11 | .64 | .57 | .46 |\n\nRESULTS AND DISCUSSION\n\nThe variations of lift, drag, pitching moment, and lift-drag ratio for an angle-of-attack range of approximately -6° to 6° are given for all wings of both the elliptical-leading-edge and wedge-leading-edge series. These characteristics at Mach numbers of 1.62, 1.92, and 2.40 may be seen in figures 5, 6, and 7, respectively, and are summarized in table 2. Similarly, the characteristics of eight flat-plate wings, with round and beveled leading edges, tested at a Mach number of 1.92, are presented in figure 8 and are summarized in table 3.\n\nLift\n\nFor the individual wings, the lift generally varies linearly with angle of attack. For this reason, the lift results can be discussed and compared with theory on the basis of lift-curve slope. It has been shown in references 4, 5, and 6, that $\\tan \\epsilon / \\tan m$ is a basic parameter in sweptback-wing or triangular-wing theory. Values of $\\tan \\epsilon / \\tan m$ greater than 1 represent a wing whose leading edge is ahead of the Mach cone, the converse being true for values of $\\tan \\epsilon / \\tan m$ less than 1. References 5, 6, and 8 have pointed out that for triangular wings with leading edges ahead of the Mach cone, the lift-curve slope has Ackeret's theoretical two-dimensional value of\n\n$$ \\left( \\frac{dC_L}{d\\alpha} \\right)_\\infty = \\frac{4}{\\sqrt{M^2 - 1}} \\quad (1) $$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:26:13.517420+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 33, "total_pages": 46, "image_filename": "19930082613_p33.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:26:15.583001+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 52, "total_pages": 114, "image_filename": "19930086061_p52.jpg", "text": "48\nNACA RM L9J07\n\nLeft semispan\nRight semispan\nUpper\nLower\n-3\n-2\n-1 P -1\n0\n0\n1\n1\n(a) $\\psi = 0^\\circ$\n\nLeft semispan\nRight semispan\nUpper\nLower\n-3\n-2\n-1 P -1\n0\n0\n1\n1\n$10^\\circ$\n(b) $\\psi = 10^\\circ$\nNACA\n\nFigure 13.- Pressure distribution about wing 1 at various angles of yaw; $\\alpha = 14.1^\\circ$.", "timestamp": "2026-07-22T06:26:23.003904+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 58, "total_pages": 98, "image_filename": "19930086073_p58.jpg", "text": "96\n\nLift coefficient, $C_L$\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n\n0 0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\delta_{a_L} = +10.8$ $\\circ$ $\\delta_{a_R} = -10.8$\nAileron deflection, $\\delta_a$, deg\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 12.—Wing plus body at 00° angle of sideslip with one aileron deflected.\n\nNACA RM A50H04", "timestamp": "2026-07-22T06:26:23.721255+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 26, "total_pages": 47, "image_filename": "19930083221_p26.jpg", "text": "24\nNACA TN No. 1824\n\n$$\n\\vec{F} = - \\int_S \\int (p-p_o) \\vec{dS} - \\int_S \\int \\rho \\vec{V} \\left[ (\\vec{V_o} + \\vec{V}) \\cdot \\vec{dS} \\right] \\quad (37)\n$$\n\nwhere vector notation is used and\n\no subscript indicating free-stream condition\n\np, $\\rho$ local static pressure and density\n\n$\\vec{V}$ local perturbation velocity vector\n\nFor the purposes of the present report, equation (37) will be modified according to the assumptions of linearized theory and the surface S restricted to a semi-infinite circular cylinder of radius r, its axis of symmetry lying along the x axis, and with one face in the x=0 plane while the other face is at x=constant. (See figure 11.)\n\nFrom linearized theory,\n\n$$\n\\frac{p}{p_o} = 1 - M_o^2 \\frac{u}{V_o}\n$$\n\nand\n\n$$\np - p_o = - \\rho_o \\left[ V_o u + \\frac{1}{2}(u^2+v^2+w^2) \\right] + \\frac{1}{2} \\rho_o M_o^2 u^2\n$$\n\n[Figure: Diagram of a semi-infinite cylinder with surfaces labeled I, II, and III. Axes x, y, z are shown. Velocity vector $V_o$ points towards surface I. Radius r is indicated.]\n\nFigure 11.- Surfaces used in study of momentum.\n\nThe end faces of the cylinder may be denoted, as in the figure, by I, II, and the curved surface by III. Then in supersonic flow, if a distribution of sources is restricted to a region downstream of I, the drag D on the body corresponding to the source distribution is given by the expression", "timestamp": "2026-07-22T06:26:25.612729+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 37, "total_pages": 44, "image_filename": "19930082566_p37.jpg", "text": "NACA TN No. 1899\n\n.8\n.6\n.4\n.2\n0\n$\\sigma_2'/\\sigma_{1t}'$\n\nR = 2\nR = 1\nR = 0.5\nR = 0\n\n0 .2 .4 .6 .8 1.0 1.2\n$\\sigma_1'/\\sigma_{1t}'$\n\n○ Failure at $1 \\times 10^6$ cycles\n△ Failure at $5 \\times 10^5$ cycles\n□ Failure at $1 \\times 10^6$ cycles\n\n[Figure: NACA logo]\n\nFigure 13.- Biaxial fatigue-stress relationship. 24S-T extruded aluminum-alloy tubing.\n\n35", "timestamp": "2026-07-22T06:26:26.331662+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 15, "total_pages": 60, "image_filename": "19930085862_p15.jpg", "text": "```markdown\nNACA RM No. L9A07\n13\n\nattack encompassed the spoilers, the spoiler effectiveness dropped off\nand the pitching-moment coefficients became approximately the same as\nfor the wing without spoilers (fig. 18). A large positive shift in the\npitching-moment curves, which occurred throughout the angle-of-attack\nrange, was obtained with the outboard spoilers on the flapped configu-\nrations as shown in figure 21. This trim change probably occurred as a\nresult of the outboard spoiler segments remaining in regions of unsepa-\nrated flow at all angles of attack. In all cases where trim changes\noccurred, larger changes were encountered with the outboard spoiler\nlocations than with the inboard locations. The magnitude of the trim\nchange was also dependent upon the spoiler projection.\n\nComparison of Aileron and Spoilers\n\nA brief comparison of the relative rolling effectiveness of the\naileron and the spoilers is presented in figure 25. The comparison is\nmade using what is considered as the optimum spoiler span and location as\ndetermined from data presented in figures 22 and 23; namely, $0.60\\frac{b}{2}$, the\ninboard end being located at $0.20\\frac{b}{2}$ and $0.375\\frac{b}{2}$ spanwise stations for the\nplain wing and the flapped configurations, respectively. It can be seen\nthat for the plain wing the rolling-moment coefficient at small angles of\nattack produced by the 0.10c projection spoilers was approximately equal\nto that which would be produced by a total aileron deflection of $25^\\circ$. At\nhigher angles of attack, however, the rolling effectiveness of the\nspoilers dropped to zero, whereas the aileron maintained considerable\neffectiveness up through the highest angle of attack investigated. For\nthe wing equipped with the extensible leading-edge flaps and the split\nflaps, the spoilers produced rolling-moment coefficients through the high\nangle-of-attack range which were equivalent to about $35^\\circ$ of total\naileron deflection. Although the rolling effectiveness of the aileron\nincreased about linearly with deflection up to $\\delta_a = \\pm 25^\\circ$, the use of\nlarge deflections for ailerons equipped with conventional internal-\nbalance systems is limited on thin wings of the type investigated herein\nto about $\\pm 15^\\circ$ for a 30-percent balance chord.\n\nThe value of the wing-tip helix angle produced in a steady roll by a\nlateral-control device is indicative of the power or effectiveness of that\ndevice. The helix angles were therefore estimated as $C_l/C_{l_p}$ where $C_{l_p}$\nis the damping-in-roll coefficient. With the aileron deflection limited\nto about $\\pm 15^\\circ$ the value of the helix angle obtainable at $0.35C_{L_{max}}$ with\nthe flaps-deflected configuration would be about 0.084 for the aileron\nand 0.093 for the 0.10c projection spoilers. This difference between the\nvalues of the helix angle produced by the aileron and by the spoiler would\nprobably be even greater if account were taken of the adverse yawing-\nmoment characteristics at the lower speeds.\n```", "timestamp": "2026-07-22T06:26:27.949622+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 24, "total_pages": 149, "image_filename": "19930083192_p24.jpg", "text": "20\nNACA TN 1976\n\nThe spacing is approximately twice the average gust-gradient distance of\n10 to 14 chords. These data indicate that the gust shape in the direc-\ntion of flight is either triangular or sinusoidal in character as\ncontrasted to the \"flat-top\" gust assumed in past years.\n\nFrom time to time, comparisons are made of the gust spacing in\nreference 11 of 22 chords with the gust spacing that was derived from\nstatistical gust data in reference 9 of 11 chords. The discrepancy be-\ntween the two figures arises from the fact that the data on gust spacing\npresented in reference 11 were obtained for the larger gusts and represent\ndata on gust intensities ranging from 5 feet per second up to the maxi-\nmum value recorded, whereas the gust spacing listed in reference 9\nis based on the average spacing of all gusts from a gust velocity of\n0.3 foot per second to the maximum value experienced by any of the air-\nplanes considered.\n\nThe sequences of gusts found in reference 11 were such that in\nabout two-thirds of the cases of sets of two the gusts were of opposite\nsign while the remaining sets of two were composed of gusts of like sign.\nThe larger gust intensity could be found at any point in the sequence\nunder consideration.\n\nLongitudinal gusts.- Inspection of the data for the maximum values\nof the horizontal and vertical gust velocities in the same traverse\n(fig. 20) indicates equality of the gust intensities in the two direc-\ntions. Figure 21 for the XBM-1 and Aeronca C-2 airplanes gives essen-\ntially the same indication, but the agreement is not exact and the\ndiscrepancy between the values of the gust velocities is an offset of\nroughly 1 foot per second. The data obtained from the Aeronca C-2 air-\nplane show the larger discrepancy. Some of this discrepancy may be due\nto the fact that the amount of horizontal-gust data was larger than the\namount of vertical-gust data for the Aeronca C-2 airplane and smaller\nthan the amount for the XBM-1 airplane.\n\nSince the maximum gust intensity in the horizontal direction is\nequal to the vertical gust intensity for the same region of turbulence\nand a preliminary investigation of frequency distributions indicates\nthat the frequency distributions are essentially the same, the atmos-\npheric turbulence seems to be isotropic. Previous indications as to\nstructure of vertical gusts, that is, gust spacing, gust-gradient\ndistance, and lateral distributions of gust velocity, would be assumed\nto apply to horizontal gusts.\n\nThe preceding results indicate equal gradient distances for the\nlongitudinal and lateral faces of the gust velocity distribution, and\nthis equality suggests that the gust is symmetrical about a vertical\naxis. In keeping with this concept, the findings suggest that the\ndistribution might be visualized as a four-sided pyramid with a base", "timestamp": "2026-07-22T06:26:29.206602+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 27, "total_pages": 37, "image_filename": "19930082646_p27.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:26:29.742397+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 3, "total_pages": 29, "image_filename": "19930085899_p3.jpg", "text": "2\nNACA RM No. L9A21\n\nThis paper presents the results of the investigation of the wing-fuselage combination employing a wing with the quarter-chord line swept back 45°, with aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil section.\n\nMODEL AND APPARATUS\n\nThe wing of the semispan model had 45° of sweepback referred to the quarter-chord line, a taper ratio of 0.60, aspect ratio of 4, and an NACA 65A006 airfoil section parallel to the free stream. The wing was made of beryllium copper and the fuselage of brass. A two-view drawing of the model is presented in figure 1 while ordinates of the fuselage of fineness ratio 10 can be found in table I. Details of the two wing fences that were used in the course of the investigation are shown in figure 2.\n\nThe model was mounted on an electrical strain-gage balance, which was enclosed in the bump, and the lift, drag, pitching moment, and bending moment about the model plane of symmetry were measured with calibrated galvanometers. The angle of attack was changed with a small electric motor and the value of the angle was determined with a calibrated slide-wire potentiometer.\n\nEffective downwash angles were determined for a range of tail heights by measuring the floating angles of five free-floating tails with the aid of calibrated slide-wire potentiometers. Details of the floating tails, which had plan forms identical to that of the wing, are shown in figure 3; while a photograph of the test set-up on the bump is given in figure 4.\n\nA total-head comb was used to determine dynamic-pressure ratios for a range of tail heights in a plane which contained the 25-percent mean-aerodynamic-chord point of the free-floating tails. The total-head tubes were spaced 0.25 inch apart.\n\nSYMBOLS\n\n$C_L$ lift coefficient $\\left(\\frac{\\text{Twice panel lift}}{qS}\\right)$\n\n$C_D$ drag coefficient $\\left(\\frac{\\text{Twice panel drag}}{qS}\\right)$\n\n$C_m$ pitching-moment coefficient referred to 0.25$\\bar{c}$ $\\left(\\frac{\\text{Twice panel pitching moment}}{qS\\bar{c}}\\right)$", "timestamp": "2026-07-22T06:26:30.037690+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 25, "total_pages": 72, "image_filename": "19930085491_p25.jpg", "text": "24 CONFIDENTIAL NACA RM No. A8J04\n\nReynolds number on the laminar separation bubble (fig. 13(c)) near the wing leading edge. Von Doenhoff and Tetervin, reference 21, have shown at subsonic speeds that increasing the Reynolds number caused a decrease in the chordwise extent of the separated bubble. This change with increased Reynolds number resulted in an increase in the negative pressures over the wing leading edge. In the present tests the less rearward inclination of the resultant force vector with increasing Reynolds number that is shown by the reduction in $k_{\\text{aopt}}$ indicates a similar increase in leading-edge suction force. This effect of Reynolds number is not realized on the outboard sections where the flow does not reattach after separating near the leading edge.\n\nThe large area of the wing near the tip where the flow does not reattach indicates the detrimental effect of the adverse pressure gradient due to angle of attack (fig. 5) which exists over these sections at lift coefficients near the optimum. The accompanying increase in pressure drag associated with this pressure gradient, particularly near the leading edge when this separation occurs, is apparent by comparing the theoretical value of $k_a$ of 0.54 for WF-63 with the values obtained near zero lift ($k_a = 0.66$) and at the optimum lift coefficient ($k_a = 0.74$) for a Reynolds number of 0.84 million. In reference 12, tests were made at the Mach number of the present study with a wing having a sharp leading edge having approximately the same length and sweep angle as WF-63 and a value of $k_a$ of 0.79 was obtained. Comparison of this experimental result with those obtained in the present investigation of WF-63 suggests that when the line of flow separation moves near the leading edge the advantage of leading edge rounding in reducing $k_a$, and consequently $\\Delta C_D / (\\Delta C_L)^2$, is apparently lost. Thus the problem of leading-edge shape with emphasis on the reduction of the strength of the adverse pressure gradient due to angle of attack should be investigated in an attempt to maintain the maximum leading-edge suction force to the highest possible lift coefficient.\n\nEffect of Reynolds number on maximum lift-drag ratio.- The experimental values of maximum lift-drag ratio obtained with WF-63 at Reynolds numbers of 0.31, 0.62, and 0.84 million were 5.8, 6.7, and 7.2, respectively. This increase, shown in figure 7(c) and in table II, with increasing Reynolds number results from the reductions in $C_{D_{\\text{min}}}$ and $\\Delta C_D / (\\Delta C_L)^2$ which were discussed in the preceding sections. Although the highest experimental value obtained with this configuration is considerably less than the theoretical value of 10.1, the favorable effect of increased Reynolds number indicates\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:26:30.520065+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 40, "total_pages": 65, "image_filename": "19930082546_p40.jpg", "text": "NACA TN No. 1870\n39\n\n[Figure: (b) Reinforced plywood wall (rear view) showing microphone support. Figure 5.- Continued.]", "timestamp": "2026-07-22T06:26:32.413527+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 23, "total_pages": 46, "image_filename": "19930085548_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:26:34.859143+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 43, "total_pages": 50, "image_filename": "19930082592_p43.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T06:26:35.284254+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 27, "total_pages": 62, "image_filename": "19930082918_p27.jpg", "text": "26\nNACA TN 1940\n\nAPPENDIX\n\nPROCESSING OF LOW-CARBON N-155 7/8-INCH BROKEN-CORNER SQUARE BAR\nSTOCK FROM HEAT A-1726\n\nThe Universal-Cyclops Steel Corporation reported the processing of\nthe low-carbon N-155 bar stock to be as follows.\n\nAn ingot was hammer cogged and then rolled to bar stock under the\nfollowing conditions:\n\n(1) Hammer cogged to a 13-inch-square billet\nFurnace temperature, $2210^\\circ$ to $2220^\\circ$ F\nThree heats - Starting temperature on die, $2050^\\circ$ to $2070^\\circ$ F\nFinish temperature on die, $1830^\\circ$ to $1870^\\circ$ F\n\n(2) Hammer cogged to a $10\\frac{3}{4}$-inch-square billet\nFurnace temperature, $2200^\\circ$ to $2220^\\circ$ F\nThree heats - Starting temperature on die, $2050^\\circ$ to $2070^\\circ$ F\nFinish temperature on die, $1790^\\circ$ to $1800^\\circ$ F\n\n(3) Hammer cogged to a 7-inch-square billet\nFurnace temperature, $2200^\\circ$ to $2220^\\circ$ F\nThree heats - Starting temperature on die, $2050^\\circ$ to $2070^\\circ$ F\nFinish temperature on die, $1790^\\circ$ to $1890^\\circ$ F\nBillets ground to remove surface defects\n\n(4) Hammer cogged to a 4-inch-square billet\nFurnace temperature, $2190^\\circ$ to $2210^\\circ$ F\nThree heats - Starting temperature on die, $2040^\\circ$ to $2060^\\circ$ F\nFinish temperature on die, $1680^\\circ$ to $1880^\\circ$ F\nBillets ground to remove surface defects\n\n(5) Hammer cogged to a 2-inch-square billet\nFurnace temperature, $2180^\\circ$ to $2210^\\circ$ F\nThree heats - Starting temperature on die, $2050^\\circ$ to $2065^\\circ$ F\nFinish temperature on die, $1730^\\circ$ to $1870^\\circ$ F\nBillets ground to remove surface defects\n\n(6) Rolled from a 2-inch-square billet to a 7/8-inch broken-corner\nsquare bar - one heat\nFurnace temperature, $2100^\\circ$ to $2110^\\circ$ F\nBar temperature at start of rolling, $2050^\\circ$ to $2060^\\circ$ F\nBar temperature at finish of rolling, $1910^\\circ$ F", "timestamp": "2026-07-22T06:26:37.961934+00:00"}