PDF/ocr_dataset_1042.jsonl

{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 40, "total_pages": 96, "image_filename": "19930085880_p40.jpg", "text": "38\nNACA RM No. L9C03\n\nLoad, lb\nSpeed\n(fps)\nWetted area, sq ft\n\n[Figure: Graph showing Load vs. Wetted area for various speeds (10, 15, 20, 25 fps).]\n\n(d) $\\tau = 16^\\circ$.\nFigure 14.- Continued.", "timestamp": "2026-07-22T05:36:35.076126+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 26, "total_pages": 31, "image_filename": "19930085881_p26.jpg", "text": "24\nNACA RM L9D12\n\nCONFIDENTIAL\n\n$C_D$\n.12\n.08\n.04\n0\n\n.16\n.12\n.08\n.04\n0\n\n$pb/2V$\n\n| Model | $\\delta_a (deg)$ | $i_w (deg)$ |\n| :--- | :---: | :---: |\n| 120b | 3.5 | 0 |\n| 120e | 4.7 | -0.13 |\n| 120f | 4.7 | 0.02 |\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\n\nNACA\n\n(b) 9-percent-thick double-wedge airfoil section.\nFigure 6.- Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:36:36.018329+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930085912_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:36:42.183718+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 12, "total_pages": 114, "image_filename": "19930086061_p12.jpg", "text": "```markdown\n8\nNACA RM L9J07\n\n# RESULTS AND DISCUSSION\n\n## CHORDWISE PRESSURE DISTRIBUTIONS AND FLOW CHARACTERISTICS\n\n### Pressure Distributions and Flow Characteristics at Zero Yaw\n\n**Presentation of data.**- The zero-yaw chordwise pressure distributions of the three wings investigated are plotted in rectangular Cartesian coordinates in figures 6 to 11 and in oblique Cartesian coordinates on isometric views of the wing plan forms in part (a) of figures 12 to 26. The zero-yaw flow characteristics as observed by surface tufts are shown in part (a) of figures 27 to 29.\n\n**Concept of correlation between pressure distributions and flow.**- Each wing semispan displayed a region of relatively high negative pressure over the upper surface that was confined to a narrow strip at the leading edge for the lowest angles of attack. With increasing angles of attack the region outboard along the semispan progressively swept back from the leading edge and inward toward the plane of symmetry; that is, the region progressively fanned out over a greater chordwise length. Immediately downstream from the high negative pressures was a lower negative-pressure region, which was well defined over the inboard stations but which spread increasingly over the outboard sections. As substantiated by smoke-flow and tuft-probing studies, these high and low negative-pressure regions were associated with conical separation vortices. The three-dimensional vortices, rotating with the bottom tangential component of velocity toward the leading edge, are illustrated schematically in figures 30(a) and 30(b) as observed over wing 2. The sections outboard along the semispan effectively operated at progressively higher resultant angles of attack and the resulting higher leading-edge negative pressures at the wing tips caused a strong spanwise flow of the low-energy boundary-layer air. Observations from a direction parallel to the wing leading edge of a narrow jet of smoke issued close to the leading edge gave a representative visual interpretation of the chordwise flow such as is sketched in figure 30(c) for wing 2 at $\\alpha = 24.1^\\circ$. The short-dash line represents a stream line at the boundary of the region of rotating turbulent flow. The boundary was distinct over the forward part of the turbulent region but became less defined farther back. The tuft probing and smoke studies indicated that the center of vortex rotation and the maximum depth of the turbulent region occurred at the chordwise position corresponding to the negative-pressure peak. Behind the point of maximum thickness, the depth of the turbulent flow diminished rather rapidly; and, as the boundary of the turbulent region bent toward the wing surface, the value of the negative pressure coefficient decreased sharply and approached more nearly the free-stream static pressure at the approximate chordwise point of contact of the boundary with the wing surface. The position of the pressure dip could be defined reliably by tufts when the vortex was strong\n```", "timestamp": "2026-07-22T05:36:42.787541+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 71, "total_pages": 99, "image_filename": "19930082511_p71.jpg", "text": "NACA TN No. 1826\n69\n\n<!-- Image (273, 123, 828, 383) -->\n\n(a) Infinitely long open tunnel.\n\n<!-- Image (229, 503, 822, 770) -->\n\n(b) Open jet between closed entrance and exit regions.\nFigure 1.— Illustrations for discussion of surface perturbation velocity\nin open wind tunnels.", "timestamp": "2026-07-22T05:36:45.075370+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 6, "total_pages": 98, "image_filename": "19930086073_p6.jpg", "text": "4\nNACA RM A9H04\n\n$\\delta_f$\nsplit-flap deflection, measured perpendicular to hinge line, degrees\n\n$\\delta_r$\nrudder deflection, measured perpendicular to hinge line, degrees\n\n$\\frac{d\\alpha_t}{d\\delta_r}$\nrate of change of angle of attack of the vertical tail with rudder deflection for constant tail normal-force coefficient\n\n$\\sigma$\nincrement of tail angle of attack above that due to the angle of sideslip, produced by sidewash at the tail, degrees\n\nEQUIPMENT\n\nThe principal dimensions of the model are given in figure 2(a) and table I. The airfoil section of the wing, taken in the streamwise direction, was developed from a symmetrical double-wedge airfoil section as described in figure 2(b). Coordinates for the body of fineness ratio 12.5 used in this investigation are presented in table II. The vertical tail had a symmetrical double-wedge airfoil section with a maximum thickness of 5-percent chord at 50-percent chord. Split-flap-type control surfaces were used on the wing, negative flap deflections being obtained by placing the flaps on the upper surface of the wing. A gap was produced in the span of the flaps by the presence of the tail boom used with the wing-alone model.\n\nThe photographs of figure 3 show the model as mounted in the Ames 40- by 80-foot wind tunnel.\n\nTESTS AND CORRECTIONS TO DATA\n\nForce and moment data were obtained through the angle-of-attack range at various angles of sideslip for the wing alone, wing plus body, and wing plus body and vertical-tail configurations as outlined in table III. The investigation was conducted at a dynamic pressure of 25 pounds per square foot, which corresponds to a Mach number of approximately 0.13 and a Reynolds number of approximately $15.4 \\times 10^6$ based on the mean aerodynamic chord.\n\nThe force and moment data are presented with reference to the stability axes with the origin located at the half-chord station of the root chord of the modified wing. The latter point corresponds to the same longitudinal station as the quarter-chord station of the mean aerodynamic chord.\n\nAll of the force data have been corrected for air-stream inclination and for wind-tunnel-wall effect, the latter correction being that for a wing of the same span having elliptic loading but with an unswept plan form. The following corrections were applied:", "timestamp": "2026-07-22T05:36:46.504340+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 3, "total_pages": 50, "image_filename": "19930086076_p3.jpg", "text": "NACA RM E9F09\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nDESIGN FACTORS FOR 4- BY 8-INCH RAM-JET COMBUSTOR\n\nBy Donald W. Male and Adolph J. Cervenka\n\nSUMMARY\n\nAn investigation was made of a series of flame holders designed with the objective of providing a high combustion efficiency in a 4- by 8-inch ram-jet combustor 24 inches long at an inlet-air velocity of 200 feet per second, inlet-air pressure of 60 inches of mercury absolute, inlet-air temperature of $200^\\circ$ F, and near stoichiometric fuel-air ratio.\n\nThe most efficient and stable combustor designs investigated employed fuel injection directly at the upstream end of a flame holder that combined surfaces heated to incandescence by immersion in flame with a continuous path of low stream velocity for flame propagation between the upstream and downstream ends of the flame holder. Silicon-coated molybdenum was satisfactorily used as flame-holder material at the high temperatures encountered.\n\nINTRODUCTION\n\nOne phase of NACA research on ram-jet combustion is the study of design principles for ram-jet combustors. A ram-jet combustor generally consists of a fuel injector followed by a fuel-air mixing length, a flame holder, and sufficient space for complete combustion.\n\nAn investigation was conducted from June 1947 to January 1949 at the NACA Lewis laboratory to study combustor design factors leading to a sufficiently high heat release and low internal drag to power a high-thrust, low-drag ram-jet engine. Two design principles investigated were the use of flame holders that employed incandescent surfaces heated by immersion in flame, as suggested in reference 1, and the elimination of the separate fuel-air mixing length by injecting the fuel directly at the upstream end of the flame holder. In this manner, combustion can exist at the point of fuel injection, allowing more efficient use of the available combustion space, and intermittent combustion in the mixing zone, a common cause of combustion instability is eliminated.", "timestamp": "2026-07-22T05:36:48.643047+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 20, "total_pages": 51, "image_filename": "19930085962_p20.jpg", "text": "```markdown\nNACA RM A9E05\n\nCONFIDENTIAL\n\n16\n14\n12\n10\n8\n6\n4\n2\n0\n-2\n-4\n-6\n-8\n\nLift coefficient, $C_L$\n\n$\\delta_e$ (deg)\n0\n2\n4\n6\n10\n20\n30\n\n-16 -12 -8 -4 0 4 8 12 16\nAngle of attack, $\\alpha$, degrees\n\n.16 .12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28 -.32\nPitching-moment coefficient, $C_m$\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 6. — The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.80.\n\nCONFIDENTIAL\n\n19\n```", "timestamp": "2026-07-22T05:36:55.217151+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 41, "total_pages": 96, "image_filename": "19930085880_p41.jpg", "text": "NACA RM No. I9C03\n39\n\n<!-- Image (138, 134, 903, 886) -->\n\n(e) $\\tau = 20^\\circ$.\nFigure 14.- Concluded.", "timestamp": "2026-07-22T05:36:56.660637+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 9, "total_pages": 33, "image_filename": "19930085977_p9.jpg", "text": "8 CONFIDENTIAL NACA RM L9H22\n\nindicated. When the fuselage was added to the isolated wing these large changes in downwash slope were not evident.\n\nThe results of the point dynamic-pressure surveys made in a vertical plane containing the 25-percent-mean-aerodynamic-chord point of the free-floating tails used in the downwash surveys are presented in figure 13. Below a Mach number of 0.95 there is very little difference in the wake characteristics of the wing-alone and wing-fuselage configurations except that larger wake losses are indicated at $\\alpha = 10^\\circ$ for the wing-fuselage condition because of a more fully developed stall. At the Mach numbers above 1.00 at moderate and high angles of attack, however, the wake associated with the wing-fuselage configuration was much more extensive than the corresponding isolated-wing wake (fig. 13). In order to gain further information concerning the possible cause of these wake differences, a few spanwise surveys were made at the same tail length used for the vertical surveys. The results of these additional surveys (fig. 14) indicated that, although the isolated-wing wake losses are practically constant along the span of the tail, a very large spanwise dynamic-pressure gradient was present near the fuselage. The flagged symbols plotted in figure 14 represent the data obtained from figure 13 for the same survey location. It is apparent from the comparison of the two sets of data that, while the wake measurements behind the wing alone could be repeated, the wake characteristics behind the wing-fuselage combination could not be repeated. These discrepancies in wake behavior as well as the steep gradient in dynamic pressure close to the wing-fuselage juncture may be attributable to unsteady flow conditions induced by shock formations and separation at the wing-fuselage juncture.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:37:00.861709+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 27, "total_pages": 31, "image_filename": "19930085881_p27.jpg", "text": "NACA RM L9D12\n25\n\nCONFIDENTIAL\n\n.12\n.08\n$C_D$\n.04\n0\n\n.16\n.12\n.08\n$pb/2V$\n.04\n0\n-.04\n\nModel $\\delta_a (deg)$ $i_w (deg)$\n--- 116a 3.5 0\n--- 116c 5.0 -0.03\n--- 116d 5.0 -0.06\n--- 116e 3.8 -0.17\n--- 116f 3.9 -0.20\n\n.6 .8 1.0 1.2 1.4 1.6 1.8 2.0\nM\nNACA\n\n(c) NACA 16-009 airfoil section.\nFigure 6.— Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:37:01.689565+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 50, "total_pages": 58, "image_filename": "19930082617_p50.jpg", "text": "NACA TN 1962\n49\n\nCY L-75\n\nFigure 34.- Bottom view of cylinder 75 after buckling.\n[NACA logo]", "timestamp": "2026-07-22T05:37:03.556863+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 60, "total_pages": 66, "image_filename": "19930082914_p60.jpg", "text": "NACA TN No. 1857\n59\n\n[Figure: A line graph plotting \"Fringe number\" on the y-axis (0 to 40) against \"Fringe position, mm from datum line\" on the x-axis (0 to 70). A straight line rises from approximately (0, 2) to (70, 37). A single data point is plotted near (43, 1). The NACA logo is in the bottom right corner of the plot area.]\n\nFigure 12.- Location of undisturbed fringes.", "timestamp": "2026-07-22T05:37:06.825818+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 23, "total_pages": 36, "image_filename": "19930085912_p23.jpg", "text": "NACA RM No. E9C16\n\nModel center line\nThermocouple rakes\nOrifices\nPressure taps\nThermocouple probes\nPlenum chamber\nFlush-type skin thermocouples\nThermocouple cross rake\nMounting strut\nAccessory housing\nStreamline support struts\nScreen\nFront pressure rakes\nStatic-pressure tubes\nRear pressure rakes\nAdjustable tail cone\n\nNACA\n\nFigure 2. - Sketch of model showing instrumentation used in investigation.\n\n21", "timestamp": "2026-07-22T05:37:07.804232+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 1, "total_pages": 28, "image_filename": "19930086092_p1.jpg", "text": "NACA RM A9F14\n\nCONFIDENTIAL\n\nCopy 271\n\nRM A9F14\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION\n\nEMPLOYING A WING SWEPT BACK 63°.- AERODYNAMIC\n\nCHARACTERISTICS IN SIDESLIP OF A LARGE-SCALE\n\nMODEL HAVING A 63° SWEPT-BACK\n\nVERTICAL TAIL\n\nBy Gerald M. McCormack\n\nAmes Aeronautical Laboratory\n\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\n\nAUTHORITY: RESEARCH ABSTRACT NO. 102\n\nDATE: JUNE 22, 1956\n\nWHL\n\nNATIONAL ADVISORY COMMITTEE\n\nFOR AERONAUTICS\n\nWASHINGTON\n\nOctober 7, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:37:10.290821+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 13, "total_pages": 114, "image_filename": "19930086061_p13.jpg", "text": "```markdown\nNACA RM L9J07\n9\n\nby establishing a narrow chordwise band where tufts at the wing surface were in a state of transition between the undisturbed rearward flow and the strong spanwise flow in the separated region of lower pressure.\n\nGenerally the same type of pressure distributions were reported in reference 8 for a two-dimensional investigation of a 6-percent-thick biconvex section; however, the pressure change following the peak negative pressure was not as great as observed in the present three-dimensional investigation. Unpublished data for the two-dimensional airfoil (investigated in reference 8) indicate a standing region of turbulent separated flow having its greatest depth at approximately the position of maximum negative pressure and decreasing in depth farther back chordwise where the reduction in negative pressure occurred. Measured velocity profiles indicated that the boundary of the separated region had the same general contour as the boundary streamline sketched in figure 30(c) for wing 2 of the present investigation. For the two-dimensional airfoil the pressure dip seemed to be just behind the chordwise location where surface tufts indicated intermittent forward and rearward flow.\n\nPressure distributions and flow characteristics of wing 2 at zero yaw.- In light of the general concepts given in the foregoing remarks, the pressure distributions and flow characteristics of wing 2, which are typical for all three wings, are discussed in detail. At an angle of attack of $4.1^\\circ$ a region of relatively high negative pressure close to the leading edge was followed by a region of lower pressure, which indicated the presence of a separation vortex along most of the semispan. The tip sections at this low angle of attack were more highly loaded than the inboard stations. One apparent reason for the higher outboard loading was the increasing induced angle of attack along the semispan such as would be expected from considerations of potential flow over a triangular wing. The areas of relatively high and low negative pressures at the short outboard chords were poorly defined (fig. 8). Two possible explanations for the characteristic decreasing chordwise pressure change outboard from station 4, caused primarily by the weakening of the negative-pressure dip behind the vortex, could be (a) an equalization of pressure throughout the thickened tip boundary layer and (b) a more gradual return of the flow above the turbulent vortex region to the wing surface. This gradual return could be caused by the vortex trying to sweep back from the leading edge as it does for higher angles of attack.\n\nWith the angle of attack increased to $8.1^\\circ$, the vortex swept back on the wing and became stronger and thus gave sharper distinction between the negative-pressure peaks and dips on the wing. The pressure distributions of figure 8 indicate that the vortex was approximately at 4, 10, 15, 30, and 65 percent of the chord of stations 2, 3, 4, 5, and 6, respectively. The vortex was increasingly hard to locate by the pressure distributions outboard from station 4 because as the vortex grew larger\n```", "timestamp": "2026-07-22T05:37:17.608763+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 12, "total_pages": 42, "image_filename": "19930085938_p12.jpg", "text": "NACA RM No. L9B04\n\n[Figure: A wind-tunnel model of an aircraft with a single boom, labeled as \"(a) With single boom. (Langley tank model 237-7B.)\"]\n\nNACA\nL-54130.1\n\nFigure 1.— Wind-tunnel models.\n\n11", "timestamp": "2026-07-22T05:37:18.335886+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 42, "total_pages": 96, "image_filename": "19930085880_p42.jpg", "text": "40\nNACA RM No. L9C03\n\nResistance, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(a) $\\tau = 4^0$.\nNACA\nFigure 15.- Variation of resistance with wetted area. Model 250A.", "timestamp": "2026-07-22T05:37:21.396745+00:00"}
{"citation_id": "19930085889", "source_url": "https://ntrs.nasa.gov/api/citations/19930085889/downloads/19930085889.pdf", "page_number": 37, "total_pages": 37, "image_filename": "19930085889_p37.jpg", "text": "36\nNACA RM L9F14\n\nCONFIDENTIAL\nSweepback, $\\Lambda$, deg\n$\\circ$ 36\n$\\square$ 32.6\n$\\diamond$ 46.7\n\n$C_{l_\\delta}$\n.002\n.001\n0\n\n$\\left(\\frac{pb}{2V}\\right)_\\delta$\n.008\n.006\n.004\n.002\n0\n\nCONFIDENTIAL\nNACA\n\n-2 0 .2 .4 .6 .8 1.0\nLift coefficient, $C_L$\n\nFigure 18.— Variation of aileron effectiveness $C_{l_\\delta}$ and $(pb/2V)_\\delta$ with lift coefficient for the wings tested with fuselage.\n\nNACA-Langley - 8-9-49 - 250", "timestamp": "2026-07-22T05:37:21.584595+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 21, "total_pages": 51, "image_filename": "19930085962_p21.jpg", "text": "20\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (78, 148, 851, 851) -->\n\n(b) $C_L$ vs $C_D$.\nFigure 6.—Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:37:22.037492+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 4, "total_pages": 50, "image_filename": "19930086076_p4.jpg", "text": "2\nNACA RM E5F09\n\nThe experimental procedure involved determining combustion stability and efficiency and isothermal pressure drop for several combustors with various flame holders and fuel injectors at a simulated flight condition. At the simulated flight condition, the combustor inlet-air conditions were: inlet-air pressure, 60 inches of mercury absolute; inlet-air temperature, 200° F; inlet-air velocity, 200 feet per second; and fuel-air ratio, 0.05.\n\nIn addition to the design problem there exists the problem of materials for flame holders, because the temperatures attained are above the melting point of readily available alloys. The use of silicon-coated molybdenum was suggested by Roger Long of the Fabrication Division, Lewis laboratory, who also assisted in the procurement of it for use in this investigation.\n\nAPPARATUS\n\nA rectangular combustor, which is adaptable to wing installations, was chosen for the investigation. A diagram of the 4- by 8-inch test installation is shown in figure 1. The combustor consisted of a constant cross-section, rectangular Inconel duct, 4 by 8 inches in cross section and 24 inches long, with an ignitor, a flame holder, and a fuel injector. The combustor and exhaust ducting were water-jacketed in order to prevent overheating and to permit measurement of heat losses. A water spray was installed at the downstream end of the combustor to cool the exhaust gases to temperatures that could be measured with thermocouples.\n\nA Lucite window in the ducting, as shown in figure 1, permitted observation of a part of the combustor and a quartz window permitted visual verification of the absence of flame downstream of the water spray.\n\nInlet diffuser. - Air flows and pressures were controlled by means of variable-area inlet-diffuser and exhaust-nozzle sections and by remotely operated butterfly valves. A needle-type variable-area inlet diffuser was used during investigation of the first 11 flame holders (fig. 1). This diffuser was 18 inches long with a 4-inch-diameter circular cross section at its inlet and a 4- by 8-inch rectangular section at its outlet. The inlet-to-outlet area ratio was varied by moving axially a conical center body or \"needle.\"\n\nIn order to have a direct view into the combustor and to take high-speed motion photographs of the combustion process, the needle-type inlet diffuser was replaced in later runs by a movable-wall diffuser. This movable-wall diffuser was rectangular in", "timestamp": "2026-07-22T05:37:22.384858+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 7, "total_pages": 98, "image_filename": "19930086073_p7.jpg", "text": "NACA RM A9H04\n5\n\n$$\n\\alpha_T = 0.719 \\ C_L\n$$\n\n$$\nC_{D_T} = 0.01255 \\ C_L^2\n$$\n\nDrag and pitching-moment tares resulting from strut interference, based on tares obtained with a rectangular wing, were applied to the data.\n\nRESULTS AND DISCUSSION\n\nThe basic results are presented in figures 4 to 19 and are summarized in figures 20 to 28.\n\nThe discontinuities which will be noted in the force and moment curves for the wing-alone model (figs. 4 to 6) correspond to those which were a characteristic of the model with the double-wedge airfoil section (reference 1). Discussion of these discontinuities and of the flow over triangular wings will be found in references 1, 2, and 3.\n\nLongitudinal Characteristics\n\nLift.— The body added to the triangular wing supported a lift equal to the lift normally carried by the wing area it covered. That this was the case can be seen by a comparison of the lift curve for the flaps-undeflected condition of the wing alone (fig. 4(a)) with the lift curve for the flaps-undeflected condition of the wing plus body (fig. 7(a)). The lift-curve slope through zero lift in each case was 0.039 per degree. The value of $C_{l_{max}}$ for the wing-alone model was 1.34. Interference of the body nose with the top of the wind-tunnel test section made it impossible to reach the angle of attack for $C_{l_{max}}$ of the wing-body model. However, the near coincidence of the two lift curves up to within $2^\\circ$ or $3^\\circ$ of the angle for maximum lift (of the wing alone) makes it appear likely that there was little or no change in the value of $C_{l_{max}}$ when the body was added.\n\nPitching moment.— The addition of the body to the wing caused only a slight forward shift of the aerodynamic-center location. The slopes of the pitching-moment curves (slopes taken over the lift-coefficient range between 0 and 0.4 in figs. 4(c) and 7(c)) indicate a shift of the aerodynamic center from 38.5 to 37.2 percent of the mean aerodynamic chord. This shift is about one-quarter the amount computed by adding body-alone data, obtained in the Ames 40- by 80-foot wind tunnel, to the wing-alone data.\n\nSplit-flap effectiveness.— In general, the flaps produced 20 percent less lift with the body on than with the body off. It is of interest to note that the decrease in flap effectiveness was in proportion to the decrease in flap area (20 percent) rather than to the", "timestamp": "2026-07-22T05:37:27.205255+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 20, "total_pages": 23, "image_filename": "19930085934_p20.jpg", "text": "NACA RM E9G12\n19\n\n1164\n\nIsentropic outlet temperature, $t_{i,2}$, °F\n\nWater-air\nratio, w/a\n\n.02\n\n.03\n\n.04\n\n.05\n\n.06\n\nNACA\n\nPressure ratio, $P_2/P_1$\n\n(b) Pressure ratio, 5.0 to 8.0.\n\nFigure 1. - Concluded. Variation of outlet temperature with pressure ratio\nfor isentropic compression. Compressor-inlet conditions: pressure, 14\ninches mercury absolute; temperature, 77° F; specific humidity, 0.", "timestamp": "2026-07-22T05:37:27.396607+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 51, "total_pages": 58, "image_filename": "19930082617_p51.jpg", "text": "50\n\nPage intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:37:28.044101+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 51, "total_pages": 78, "image_filename": "19930082618_p51.jpg", "text": "```markdown\nNACA TN 1945\n\n[Graph: Section lift coefficient, $c_l$, vs. Section angle of attack, $\\alpha_{0L}$, deg, and Moment coefficient, $c_{m_{c/4}}$, vs. Section angle of attack, $\\alpha_{0L}$, deg]\n\nR\n$\\circ$ 0.7 x $10^6$\n$\\diamond$ 1.0\n$\\triangle$ 2.0\n$\\square$ 4.0\n$\\triangledown$ 6.0\nFlagged symbols denote\nstandard roughness\n\n[NACA logo]\n\n(b) Section lift and pitching-moment characteristics of the NACA 66$_2$-415 airfoil section with a 0.20c simulated split flap deflected 60$^\\circ$.\n\nFigure 10.- Continued.\n\n64\n```", "timestamp": "2026-07-22T05:37:31.062099+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 48, "total_pages": 53, "image_filename": "19930082542_p48.jpg", "text": "Page intentionally left blank\n\nPage intentionally left blank", "timestamp": "2026-07-22T05:37:31.282128+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 24, "total_pages": 36, "image_filename": "19930085912_p24.jpg", "text": "22\nNACA RM No. E9C16\n\n[Figure: A circular diagram showing calculated jet outlines. The diagram includes various dimensions and angles. Key labels include:\n- \"3/4\" diam.\" at the top center.\n- \"13/32\"\" at multiple points around the circle.\n- \"1/2\"\" at multiple points around the circle.\n- \"110°\" indicating angles between certain lines.\n- \"Model center line\" pointing to the central axis of the diagram.\n- \"NACA\" logo at the bottom right of the diagram.]\n\nFigure 3. - Sketch showing calculated jet outlines at accessory-housing tip using orifice configuration at inlet consisting of three 3/4-inch, three 1/2-inch, and six 13/32-inch orifices.", "timestamp": "2026-07-22T05:37:33.496720+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 13, "total_pages": 42, "image_filename": "19930085938_p13.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:37:41.212578+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 2, "total_pages": 28, "image_filename": "19930086092_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:37:45.892512+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 43, "total_pages": 96, "image_filename": "19930085880_p43.jpg", "text": "NACA RM No. L9C03\n41\n\n[Figure: A line graph plotting Resistance (lb) on the y-axis against Wetted area (sq ft) on the x-axis. The y-axis ranges from 0 to 8. The x-axis ranges from 0 to .35. There are five data series plotted, each representing a different speed in fps, indicated by labels on the right side of the graph: 10, 15, 20, 25, and 30. The data points are marked with different symbols (circles, squares, diamonds, triangles, inverted triangles). In the top left corner of the graph area, there is a small rectangular box with a horizontal line below it.]\n\nWetted area, sq ft\n(b) $\\tau = 8^\\circ$.\nFigure 15.- Continued.\nNACA", "timestamp": "2026-07-22T05:37:49.090223+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 72, "total_pages": 99, "image_filename": "19930082511_p72.jpg", "text": "70\nNACA TN No. 1826\n\n[Figure: Diagram of a contracting jet with an arrow indicating flow direction]\n\n(a) Contracting jet. The pressure on the free surface exceeds\nboth the upstream and downstream pressure, but is very\nclose to the upstream pressure.\n\n[Figure: Diagram of an expanding jet with an arrow indicating flow direction]\n\nNACA\n\n(b) Expanding jet. The pressure on the free surface is less\nthan either the upstream or downstream pressure, but is\nvery close to the upstream pressure.\n\nFigure 2.- Contracting and expanding jets (2 or 3 dimensions).", "timestamp": "2026-07-22T05:37:50.084374+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 61, "total_pages": 66, "image_filename": "19930082914_p61.jpg", "text": "60\nNACA TN No. 1857\n\nTop view of\nnozzle\n\nSide view of\nnozzle\n\n10\"\n\nFigure 13.- Schematic diagram of nozzle, jet, and mixing region.", "timestamp": "2026-07-22T05:37:50.647357+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 14, "total_pages": 114, "image_filename": "19930086061_p14.jpg", "text": "10\nNACA RM L9J07\n\nand as the boundary layer thickened, a less defined peak-negative-\npressure region resulted. The pressure distributions indicate station 7\nto be stalled, apparently from leading-edge separation since the vortex\nwas behind station 7 at $\\alpha = 8.1^\\circ$. Surface tufts, however, showed only\nthe usual strong spanwise flow with no visible indication of stall\n(fig. 28(a)). This same characteristic tuft behavior was noted at all\nangles of attack for the wing area ahead of the vortex.\n\nIncreasing the angle of attack to $14.1^\\circ$ and to $24.1^\\circ$ continued the\ntrends of increasing the vortex size and strength and of sweeping it\nback farther from the wing leading edge and inward toward the plane of\nsymmetry. The separation vortex at $\\alpha = 24.1^\\circ$ caused a negative-\npressure peak at the center station 1. The difference in the pressure\ncoefficient from -3.0 to -0.4 between the 10-percent and 30-percent\nchordwise orifices of station 2 ($0.167 \\text{ b/2}$) indicates that the vortex\nwas very strong. A pressure coefficient of -2.1 was measured at the 20-\npercent chord of station 3 in the peak-negative-pressure region, but at\nstation 4 the vortex was relatively parallel to the air stream and too\nlarge to influence the attainment of an outstanding peak-negative-pressure\ncoefficient. At $\\alpha = 24.1^\\circ$ the vortex at the tip swept inward enough\ntoward the plane of symmetry so as not to be behind stations 6 and 7. As\nexpected, figure 8(b) shows that these two stations remained stalled;\nhowever, the negative pressures on the upper surface were considerably\nincreased over those at $\\alpha = 14.1^\\circ$, with the net result that the stations\ndeveloped more lift at $\\alpha = 24.1^\\circ$. At $\\alpha = 24.1^\\circ$, but not at $14.1^\\circ$,\nsurface tufts at stations 6 and 7 indicated decisive stall (fig. 28(a))\nsuch as noted for the wing tips of the original DM-1 glider configuration\nwith rounded nose in reference 1 and the round-nose delta wings of refer-\nence 7, all of which had trailing-edge separation.\n\nThe pressure distributions, smoke-flow observations, and tuft\nstudies showed that further increases in the angle of attack merely con-\ntinued the trends of increasing the size and sweepback of the vortex and\nof increasing the area of outboard stall until at $\\alpha = 44.1^\\circ$ practically\nthe complete wing was stalled. The progression of the regions of rela-\ntively high and low negative pressure over wing 2 is shown very\neffectively by the pressure distributions plotted over an isometric view\nof the wing in part (a) of figures 17 to 21.\n\nThe pressure distributions, as previously described for a Reynolds\nnumber of about $0.85 \\times 10^6$, were essentially unaffected by increasing the\nReynolds number to $1.42 \\times 10^6$. The change in vortex location was\nnegligible as indicated by the pressures; however, the peak pressure\ncoefficients were generally higher and small changes in the area of the\npressure-distribution curves occurred with the increased Reynolds\nnumber.", "timestamp": "2026-07-22T05:37:56.521010+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 10, "total_pages": 33, "image_filename": "19930085977_p10.jpg", "text": "NACA RM L9H22 CONFIDENTIAL 9\n\nREFERENCES\n\n1. Weil, Joseph, and Goodson, Kenneth W.: Aerodynamic Characteristics of a Wing with Quarter-Chord Line Swept Back 45°, Aspect Ratio 4, Taper Ratio 0.6, and NACA 65A006 Airfoil Section. Transonic-Bump Method. NACA RM L9A21, 1949.\n\n2. Sleeman, William C., Jr., and Becht, Robert E.: Aerodynamic Characteristics of a Wing with Quarter-Chord Line Swept Back 35°, Aspect Ratio 4, Taper Ratio 0.6, and NACA 65A006 Airfoil Section. Transonic-Bump Method. NACA RM L9B25, 1949.\n\n3. King, Thomas J., Jr., and Myers, Boyd C., II: Aerodynamic Characteristics of a Wing with Quarter-Chord Line Swept Back 60°, Aspect Ratio 4, Taper Ratio 0.6, and NACA 65A006 Airfoil Section. Transonic-Bump Method. NACA RM L9C27, 1949.\n\n4. Schneiter, Leslie E., and Ziff, Howard L.: Preliminary Investigation of Spoiler Lateral Control on a 42° Sweptback Wing at Transonic Speeds. NACA RM L7F19, 1947.\n\n5. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947.\n\n6. Multhopp, H: Aerodynamics of the Fuselage. NACA TM 1036, 1942.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:37:59.957294+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 22, "total_pages": 51, "image_filename": "19930085962_p22.jpg", "text": "NACA RM A9E05\nCONFIDENTIAL\n\nLift coefficient, $C_L$\n14\n12\n10\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n-.8\n-1.0\n\n$\\delta_e$\n(deg)\n$\\circ$ 0\n$\\square$ 2\n$\\diamond$ 4\n$\\triangle$ 6\n$\\nabla$ 10\n$\\blacktriangle$ 20\n$\\blacktriangledown$ 30\n\nAngle of attack, $\\alpha$, deg\n-16 -12 -8 -4 0 4 8 12 16\n\nPitching-moment coefficient, $C_m$\n.16 .12 .08 .04 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28 -.32\n\n(a) $C_L$ vs $\\alpha$, $C_L$ vs $C_m$.\n\nFigure 7.—The effect of elevator deflection on the aerodynamic characteristics of the tail at a Mach number of 0.85.\n\nCONFIDENTIAL\nNACA\n21", "timestamp": "2026-07-22T05:38:00.159031+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 28, "total_pages": 31, "image_filename": "19930085881_p28.jpg", "text": "26\nNACA RM L9D12\n\nCONFIDENTIAL\n\n<!-- Image (128, 199, 768, 410) -->\n\n<!-- Image (128, 508, 852, 828) -->\n\n(a) 9-percent-thick circular-arc airfoil section.\nFigure 6.- Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:38:03.247949+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 52, "total_pages": 58, "image_filename": "19930082617_p52.jpg", "text": "NACA TN 1962\n\n51\n\n[Figure: Side view of a buckled cylindrical structure labeled \"CYL 76\"]\n\nFigure 35.- Side view of cylinder 76 after buckling.", "timestamp": "2026-07-22T05:38:05.748770+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 11, "total_pages": 46, "image_filename": "19930085972_p11.jpg", "text": "NACA RM L9B18\n9\n\nTABLE I\n\nPHYSICAL CHARACTERISTICS OF THE VARIABLE-SWEEP MODEL\n\nCenter of gravity, all sweep angles, percent chord ($c^t$ at $\\Lambda = 0^\\circ$) . 25\n\nWing:\nRoot and tip sections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T05:38:11.241285+00:00"}
{"citation_id": "19930085934", "source_url": "https://ntrs.nasa.gov/api/citations/19930085934/downloads/19930085934.pdf", "page_number": 21, "total_pages": 23, "image_filename": "19930085934_p21.jpg", "text": "20\nNACA RM E9G12\n\n[Figure: A graph plotting Isentropic enthalpy change against Pressure ratio. The graph contains multiple curves representing different water-air ratios.]\n\nIsentropic enthalpy change, $\\Delta h_{m,a}$, Btu/lb\nWater-air ratio, w/a\n0\n.02, .03\n.04 to .05\n\nPressure ratio, $P_2/P_1$\n(a) Pressure ratio, 1.5 to 5.0.\n\nNACA\n\nFigure 2. - Variation of enthalpy change with pressure ratio for isentropic compression.\nCompressor-inlet conditions: pressure, 14 inches mercury absolute; temperature, 77° F;\nspecific humidity, 0.\n\n1164", "timestamp": "2026-07-22T05:38:11.459254+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 8, "total_pages": 98, "image_filename": "19930086073_p8.jpg", "text": "6\nNACA RM A9H04\n\ndecrease in wing area influenced by the flap (a 32-percent decrease). It has been noted also, from the data in reference 4, that a decrease occurred in flap lift effectiveness proportional to the decrease in flap area for a plain flap on a wing of triangular plan form.\n\nIt is believed that there was no carry-over of flap lift effectiveness across the body. The incremental span load distribution due to deflecting a plain flap was found, from an investigation conducted in the Ames 40- by 80-foot wind tunnel, to be nearly elliptic in form for a wing-alone model. The portion of the loading for that region of the wing which would be occupied by the body was removed from the loading diagram. The change in the load was found to be very nearly 20 percent of the total load. This agrees with the change found by the force test reported herein and would indicate little or no carry-over of lift due to flap deflection.\n\nIt should also be noted that, as reported in reference 1, the variation of lift with flap deflection was nonlinear, and the variation was found to be dependent upon the flow conditions over the wing; for example, whether the angle of attack was above or below the angle for the break in the lift curve. (In fig. 20, the angles of attack of $0^\\circ$ and $8^\\circ$ represent values below the break while the angles of $19^\\circ$ and $24^\\circ$ represent those above.)\n\nIt appears from the lift curves of figure 4(a) that split flaps are of little or no value as a means of increasing $C_{l_{max}}$. Large flap deflections resulted in a reduced value of $C_{l_{max}}$.\n\nThe pitching-moment effectiveness of the flaps is represented in figure 21 by the increments of pitching-moment coefficient due to a given flap deflection for the same angles of attack at which the lift increments were presented in figure 20. The change in the increment of pitching moment, due to the addition of the body, was also found to be nearly proportional to the change in flap area.\n\nLift-drag ratios.— Adding the body to the wing reduced the $(L/D)_{max}$ value from 11 to 8.5. (See fig. 22(a).) At the same time the lift coefficient for $(L/D)_{max}$ was raised from 0.15 to 0.20. Both these effects would be expected due to the added drag of the body. The effect of flap deflection on the L/D values for the wing plus body model is presented in figure 22(b). With the controls deflected up, as needed for trim, there was an appreciable loss in L/D throughout the entire lift range.\n\nLateral and Directional Characteristics\n\nLateral and directional stability.— The stability derivatives presented in figure 23 represent the slope through zero angle of sideslip of the curves of $C_l$, $C_n$, and $C_Y$ as functions of $\\beta$ ($C_L$ constant). It", "timestamp": "2026-07-22T05:38:12.668534+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 5, "total_pages": 50, "image_filename": "19930086076_p5.jpg", "text": "NACA RM E9F09\n\ncross section and was fabricated in two parts. The upstream part of the diffuser consisted of two movable walls mounted between two parallel fixed walls 8 inches long, 4 inches wide, and 2 inches apart. The two movable walls were hinged 4 inches apart at the downstream end and were remotely powered to move symmetrically. The downstream part of the diffuser had a fixed transition section 18 inches long diffusing from a 2- by 4-inch area to a 4- by 8-inch area.\n\nA typical velocity profile at the exit of the movable-wall diffuser was lopsided. The profile was made symmetrical by the addition of a thin plate in the center of the diffuser, as indicated in figure 1. With both diffusers the velocity profiles indicated that the exit velocity was low in the center and high near the sides.\n\nExhaust nozzle. - A two-dimensional variable-area exhaust nozzle was used in the investigation of flame holders 1 to 9. The area was controlled by axial movement of a wedge into the throat section.\n\nFuel injectors. - Two types of fuel injector (fig. 2) were used. The first type consisted of tubes running across the duct parallel to the axis of the flame-holder components. Each tube had two rows of seven orifices for injection of the fuel and was designed to give a fuel-pressure drop of about 10 pounds per square inch at the lower fuel rates used in the investigations. This fuel injector, in a position directly upstream of the flame holder, was used in all the investigation unless otherwise stated.\n\nThe second type of fuel injector consisted of nine 30-gallon-per-hour, $80^{\\circ}$ hollow-cone, pressure-atomizing spray nozzles (rated at 100 lb/sq in. pressure differential) arranged to spray the fuel in the downstream direction. With both types of fuel injector, the fuel-injection ports were located at the upstream end of the flame holder and injected fuel in the downstream direction directly in line with the flame-holding bodies.\n\nFuels. - The fuels used in the investigation were liquid gasoline, specification AN-F-48b, liquid isopentane, and gaseous isopentane.\n\nIgnitor. - An ignitor consisting of a high voltage electrode and a small fuel line shielded by a small cone with the spark arcing between the electrode and the tip of the fuel line was used for starting. The ignitor spark and ignitor fuel were always turned off during the runs.", "timestamp": "2026-07-22T05:38:12.922846+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 21, "total_pages": 42, "image_filename": "19930085914_p21.jpg", "text": "20\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:38:14.112566+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 14, "total_pages": 42, "image_filename": "19930085938_p14.jpg", "text": "NACA RM No. L9B04\n\n[Figure: A model of an aircraft with twin booms.]\n\n(b) With twin booms. (Langley tank model 237-7TB.)\n\nFigure 1.- Continued.\n\nNACA\nL-58174\n\n13", "timestamp": "2026-07-22T05:38:15.195833+00:00"}
{"citation_id": "19930085912", "source_url": "https://ntrs.nasa.gov/api/citations/19930085912/downloads/19930085912.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930085912_p25.jpg", "text": "NACA RM No. E9C16\n23\n\n[Figure: Two plots showing pressure distribution around an airfoil at different angles of attack. The top plot shows data points for angles of attack 0, 4, and 8 degrees. The bottom plot shows similar data with a vertical axis labeled \"Pressure coefficient, S\" ranging from 0.3 to 1.5.]\n\nAngle of attack (deg)\n0\n4\n8\n\nPressure coefficient, S\n1.5\n1.2\n.9\n.6\n.3\n\nNACA\n\nFigure 4. - Effect of angle of attack on lip-pressure distribution without bleedback.", "timestamp": "2026-07-22T05:38:19.554206+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 3, "total_pages": 28, "image_filename": "19930086092_p3.jpg", "text": "NACA RM A9F14 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING SWEPT BACK $63^\\circ$.— AERODYNAMIC CHARACTERISTICS IN SIDESLIP OF A LARGE-SCALE MODEL HAVING A $63^\\circ$ SWEPT-BACK VERTICAL TAIL\n\nBy Gerald M. McCormack\n\nSUMMARY\n\nAn investigation has been conducted to determine the effects of a vertical tail having the leading edge swept back $63^\\circ$ on the aerodynamic characteristics of a wing-fuselage combination employing a wing with the leading edge swept back $63^\\circ$. The aerodynamic characteristics in sideslip with and without the vertical tail are presented. Included also are the rudder effectiveness and the rudder hinge-moment characteristics.\n\nAt angles of attack from $0^\\circ$ to $12^\\circ$, the effectiveness of the vertical tail was maintained to an angle of sideslip of $25^\\circ$ (the highest tested). At an angle of attack of $21^\\circ$, however, effectiveness was maintained only to an angle of sideslip of about $7^\\circ$; beyond $7^\\circ$ the directional stability was irregular.\n\nThe rudder was effective throughout the range of angles of attack and of angles of sideslip tested. At an angle of attack of $21^\\circ$, however, at angles of sideslip greater than about $9^\\circ$, rudder effectiveness was considerably less than at the lower angles.\n\nThe experimental characteristics are compared with characteristics computed theoretically in order to provide a basis for estimating the effects of geometric changes of the vertical tail.\n\nINTRODUCTION\n\nA possible aircraft configuration suitable for flight at Mach numbers up to about 1.5 is undergoing study in the research facilities of the Ames Aeronautical Laboratory to provide information relative to its aerodynamic behavior over a range of Reynolds numbers and Mach numbers. The configuration is based on the principles outlined in reference 1 which indicate that aircraft employing highly swept slender wings should be capable of relatively efficient flight ($L/D \\approx 10$) at moderate supersonic speeds. The\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:38:23.066608+00:00"}
{"citation_id": "19930085881", "source_url": "https://ntrs.nasa.gov/api/citations/19930085881/downloads/19930085881.pdf", "page_number": 29, "total_pages": 31, "image_filename": "19930085881_p29.jpg", "text": "NACA RM L9D12\n27\n\nCONFIDENTIAL\n\nNACA 65A009\n$\\phi = 10.6^\\circ$\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 50d | $52^\\circ$ |\n| 50e | $58^\\circ$ |\n| 50f | $60^\\circ$ |\n\nDouble-wedge\n$\\phi = 10.2^\\circ$\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 119b | $50^\\circ$ |\n\n$\\frac{C_{l\\delta}}{C_l}$\n\nNACA 16-009\n$\\phi = 21.0^\\circ$\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 115e | $40^\\circ$ |\n| 115f | $38^\\circ$ |\n\nCircular-arc\n$\\phi = 20.4^\\circ$\n\n| Model number | $\\delta_a$ |\n| :--- | :--- |\n| 117c | 50 |\n| 117d | 51 |\n\nM\n\nNACA\n\n(a) $\\Lambda = 0^\\circ$.\n\nFigure 7.— Summary of wing-aileron rolling effectiveness results.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:38:32.698486+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 11, "total_pages": 33, "image_filename": "19930085977_p11.jpg", "text": "10\nNACA RM L9H22\n\nCONFIDENTIAL\n\nTABLE I.- FUSELAGE ORDINATES\n\n[Basic fineness ratio 12; actual fineness ratio 10\nachieved by cutting off the rear one-sixth of\nthe body; $\\overline{c}/4$ located at $l/2$]\n\n[Figure: Diagram of fuselage shape with dimensions $l = 14.14''$, $l/2$, $5/6 l$, $x$, $D(Max)$]\n\n| Ordinates | | | |\n| :--- | :--- | :--- | :--- |\n| $x/l$ | $r/l$ | $x/l$ | $r/l$ |\n| 0 | 0 | .4500 | .04143 |\n| .005 | .00231 | .5000 | .04167 |\n| .0075 | .00298 | .5500 | .04130 |\n| .0125 | .00428 | .6000 | .04024 |\n| .0250 | .00722 | .6500 | .03842 |\n| .0500 | .01205 | .7000 | .03562 |\n| .0750 | .01613 | .7500 | .03128 |\n| .1000 | .01971 | .8000 | .02526 |\n| .1500 | .02593 | .8338 | .02000 |\n| .2000 | .03090 | .8500 | .01852 |\n| .2500 | .03465 | .9000 | .01125 |\n| .3000 | .03741 | .9500 | .00439 |\n| .3500 | .03933 | 1.0000 | 0 |\n| .4000 | .04063 | | |\n\nL. E. radius = 0.00052\n\nNACA\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:38:33.777141+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 23, "total_pages": 51, "image_filename": "19930085962_p23.jpg", "text": "22\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (128, 109, 826, 878) -->\n\n(b) $C_L$ vs $C_D$.\nFigure 7. — Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:38:35.137822+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 1, "total_pages": 36, "image_filename": "19930086097_p1.jpg", "text": "UNCLASSIFIED\nCopy 24\nCONFIDENTIAL\nRM A9H11\n\nNACA RM A9H11\n\nNACA\nGROUP 4\nDowngraded at 3 year\nintervals; declassified\nafter 12 years\n\nRESEARCH MEMORANDUM\n\nREDUCTION OF PROFILE DRAG AT SUPERSONIC VELOCITIES\n\nBY THE USE OF AIRFOIL SECTIONS HAVING\n\nA BLUNT TRAILING EDGE\n\nBy Dean R. Chapman\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCASE FILE\nCOPY\n\nJPL LIBRARY\nCALIFORNIA INSTITUTE OF TECHNOLOGY\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information\naffecting the National Defense of the United\nStates within the meaning of the Espionage Act,\nUSC 31 and 32. Its transmission or the\nrevelation of its contents in any manner to an\nunauthorized person is prohibited by law.\nInformation so classified may be imparted\nonly to persons in the military and naval\nservices of the United States, appropriate\ncivilian officers and employees of the Federal\nGovernment who have a legitimate interest\ntherein, and to United States citizens of known\nloyalty and discretion who of necessity must be\ninformed thereof.\n\nClassification Changed to\nUNCLASSIFIED\nAuthority\nDec 20a 5-200.10\nDate\n10/31/69\nBy\nS. Redwood\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nNovember 1, 1949\n\nCONFIDENTIAL\nUNCLASSIFIED\nNOV 7 1949", "timestamp": "2026-07-22T05:38:38.187744+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 44, "total_pages": 96, "image_filename": "19930085880_p44.jpg", "text": "42\nNACA RM No. L9C03\n\nResistance, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\n0\n.05\n.10\n.15\n.20\n.25\n.30\n.35\nWetted area, sq ft\n(c) $\\tau = 12^\\circ$.\nFigure 15.- Continued.", "timestamp": "2026-07-22T05:38:38.574647+00:00"}