PDF/ocr_dataset_1000.jsonl

{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 6, "total_pages": 34, "image_filename": "19930086022_p6.jpg", "text": "4\nNACA RM LGE24\n\n$M_a$\ntwice moment area of aileron measured behind aileron hinge line, cubic feet $\\left( \\int_{0.5b/2}^{0.975b/2} c'_a{}^2 dy \\right)$\n\nS\nwing area, square feet\n\n$S_a$\naileron area behind hinge line, square feet\n\n$\\overline{c}$\nwing mean aerodynamic chord measured parallel to plane of symmetry $\\left( \\frac{2}{S} \\int_0^{b/2} c^2 dy \\right)$\n\n$c_a$\nroot-mean-square chord of aileron measured normal to aileron hinge line, feet\n\n$c'_a$\nlocal aileron chord measured perpendicular to aileron hinge line\n\nc\nlocal wing chord measured parallel to plane of symmetry, feet\n\nc'\nlocal wing chord measured perpendicular to 0.273 chord line, feet\n\n$\\overline{c}_b$\nroot-mean-square chord of hypothetical aileron balance measured ahead of and normal to aileron hinge line, feet\n\n$\\Lambda$\nsweep angle of wing leading edge, degrees\n\ny\nspanwise coordinate, feet\n\n$\\alpha$\nangle of attack, degrees\n\n$\\delta_a$\naileron deflection, measured perpendicular to the aileron hinge line, degrees\n\n$\\delta_{a_{total}}$\narithmetical sum of equal up and down aileron deflections for an assumed set of ailerons\n\n$C_{l_\\delta}$\nrate of change of rolling-moment coefficient with aileron deflection\n\n$C_{h_\\delta}$\nrate of change of aileron hinge-moment coefficient with aileron deflection", "timestamp": "2026-07-22T04:08:06.601697+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 57, "total_pages": 65, "image_filename": "19930085843_p57.jpg", "text": "```markdown\nNACA RM L9C31\n55\n\n<!-- Image (113, 129, 909, 840) -->\n\nFigure 16.- A comparison of drag coefficient through the Mach number range as measured by the semispan and sting test methods on models of a tailless airplane. Vertical fins off; $\\delta_a = 0^\\circ$.\n```", "timestamp": "2026-07-22T04:08:10.180767+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 32, "total_pages": 47, "image_filename": "19930085918_p32.jpg", "text": "NACA RM A9D29\n31\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nPressure coefficient, P\n-1.6\n-1.2\n-.8\n-.4\n0\n.4\n.8\n\n-1.2\n-.8\n-.4\n0\n.4\n.8\n\n-.8\n-.4\n0\n.4\n.8\n\nSpanwise\nstation, 2y/b\n28.1%\n57.4%\n85.0%\n\nChordwise station, x/c\n.2 4 6 8 10\n.2 4 6 8 10\n.2 4 6 8 10\n\nNACA\n\n(a) $\\alpha=0.1^\\circ$\n\nFigure 9.—Chordwise pressure distributions for 45° swept-\nforward wing with the inboard half span of leading-edge\nflap deflected 30° down and the outboard half span\ndeflected 10° up.", "timestamp": "2026-07-22T04:08:12.748286+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 10, "total_pages": 54, "image_filename": "19930086015_p10.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 9\n\nis small. The effect on the characteristics of other models needs to be investigated.\n\n2. The effects of small yaw angles on the characteristics. These are, of course, small if the characteristics studied are not greatly influenced by the angle of yaw. In this regard, if the model is tested with the plane of the wing vertical, stream angularity influences only those characteristics which are functions of both pitch and yaw which is the unusual case, but which might be noted as applying to the rolling moment and yawing moment due to sideslip for swept wings and to certain characteristics of cruciform wings. Since the lift, drag, and pitching moment do not vary appreciably with small angles of yaw, these characteristics are not affected.\n\n3. The effects of axial pressure gradients in providing a buoyant force and the possibility of the pressure gradients altering the true pressure gradients over wing and body to such an extent that the viscosity effects are changed. This latter effect is remote, however (unless the stream contains discrete shock waves not revealed by the surveys). A correction may be applied for the former. (See reference 3.)\n\nIt should be noted that the results of the force tests indicate that experimental investigations of the loading due to angle of attack through measurement of the pressure difference between the upper and lower surface of a wing may as well be done with the model mounted with wing horizontal if more convenient. This may be deduced from the fact that the lift-curve and moment-curve slopes are not influenced by the orientation of the model. Tests were made with the model of reference 2 to demonstrate the validity of this conclusion. The results are shown in figure 14 for the wing horizontal and wing vertical.⁵ The agreement is seen to be generally satisfactory except in regions where viscosity effects are large near the trailing edge and tip. In these regions the pressures vary to some extent from test to test with the same model orientation.\n\n---\n\n⁵The data given are for $3.70^\\circ$ and $5.74^\\circ$ change in angle of attack with the model horizontal. Data for a nominal change in angle of $5^\\circ$ with wing vertical were obtained by rotating the $5^\\circ$ bent sting used in the tests of reference 2 through $90^\\circ$. The change in angle of attack $\\Delta\\alpha$ has been corrected for the deflection of the model support sting under load.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:08:12.993882+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 28, "total_pages": 59, "image_filename": "19930085936_p28.jpg", "text": "NACA RM No. E9B03\n27\n\n<!-- Image (152, 123, 854, 867) -->\n\n(a) $\\theta = 0^\\circ$ longitudinal plane.\nFigure 5. - Pressure distributions along longitudinal planes at $0^\\circ$ yaw angle for range of angles of attack.", "timestamp": "2026-07-22T04:08:16.870150+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 39, "total_pages": 92, "image_filename": "19930085930_p39.jpg", "text": "38\nNACA RM L9G07\n\n[Figure: Diagram showing a curved flow field with expansion and compression waves. Labels include: $15^\\circ$, $M = 1.71$, Expansion (dashed line), Compression (solid line). Stamps: UNCLASSIFIED, CONFIDENTIAL. NACA logo at bottom right.]\n\n(a) No boundary-layer separation taken into consideration.\nFigure 10.- Characteristics net for model 4.", "timestamp": "2026-07-22T04:08:17.203445+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 89, "total_pages": 122, "image_filename": "19930082090_p89.jpg", "text": "NACA TN No. 1455\n87\n\nVentilating air\n5.0\"\nI.D.\n30\"\n28\"\n7.83\"\nI.D.\nExhaust\ngas\n15\"\n3.25\"\n2.25\"\n24.5\"\n7.83\"\nI.D.\nExhaust\ngas\n56.5\"\n54\"\n5.0\"\nI.D.\no Static-pressure tap\nx Temperature traverse\nNACA\nVentilating air\n\nFigure 42.- Schematic diagram of test setup of heat exchanger K and air\nshroud, showing location of static-pressure and temperature measuring\nstations.", "timestamp": "2026-07-22T04:08:18.554314+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 11, "total_pages": 36, "image_filename": "19930086003_p11.jpg", "text": "NACA RM L9I08 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 an NACA 65A006 Airfoil Section. Transonic-Bump Method. NACA RM L9A21, 1949.\n\n2. Myers, Boyd C., II, and King, Thomas J., Jr.: Aerodynamic Characteristics of a Wing with Quarter-Chord Line Swept Back 45°, Aspect Ratio 4, Taper Ratio 0.3, and NACA 65A006 Airfoil Section. Transonic-Bump Method. NACA RM L9E25, 1949.\n\n3. 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\n4. DeYoung, John: Theoretical Additional Span Loading Characteristics of Wings with Arbitrary Sweep, Aspect Ratio, and Taper Ratio. NACA TN 1491, 1947.\n\n5. 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 L9G27, 1949.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:08:19.878673+00:00"}
{"citation_id": "19930085957", "source_url": "https://ntrs.nasa.gov/api/citations/19930085957/downloads/19930085957.pdf", "page_number": 23, "total_pages": 27, "image_filename": "19930085957_p23.jpg", "text": "```markdown\n22\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\n-4 0 4 8 12 16 20 24 28\n$\\alpha$, deg\n\n0 .04 .08 .12 .16 .20 .24 .28\n$C_D$\n\n.04 0 -.04 -.08 -.12 -.16\n$C_m$\n\n[Graph: The image displays three aerodynamic characteristic plots. The left plot shows $C_L$ vs $\\alpha$, the middle plot shows $C_L$ vs $C_D$, and the right plot shows $C_L$ vs $C_m$. Two curves are plotted in each graph, labeled \"Smooth\" and \"Rough\". The \"Smooth\" curve generally shows higher lift coefficients than the \"Rough\" curve. A NACA logo is visible in the bottom right corner of the graph area.]\n\n(a) Split flaps off; $\\delta_F = 60^\\circ$; $R = 4.7 \\times 10^6$.\n\nFigure 11.- Aerodynamic characteristics of a $42^\\circ$ sweptback wing with $0.575 \\frac{b}{2}$ leading-edge flaps with and without leading-edge roughness.\n\nNACA RM L59E02\n```", "timestamp": "2026-07-22T04:08:20.482378+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 56, "total_pages": 118, "image_filename": "19930085838_p56.jpg", "text": "54\nNACA RM No. L9B23\n\n<!-- Image (113, 109, 812, 937) -->\n\nAileron section hinge-moment coefficients, $C_{h_a}$\n\nSection angle of attack, $\\alpha_{02}$ deg\n(k) $\\delta_f = 40^\\circ$.\nFigure 7.- Continued.", "timestamp": "2026-07-22T04:08:21.061630+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 25, "total_pages": 50, "image_filename": "19930085952_p25.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:08:26.799895+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 55, "total_pages": 82, "image_filename": "19930085551_p55.jpg", "text": "54\nNACA RM No. L8K30\n\nYawing velocity,\nLeft rad/sec Right\nPitching velocity,\nDown rad/sec Up\nControl force,\nlb Pull Right\nPush Left\nControl position,\ndeg Right Up\nLeft Down\nAltitude,\nft\nIndicated\nairspeed,\nmph\n\nTime, sec\n\n(e) Approach condition; flaps 20°; gear down; power for level flight;\nno attempt to control altitude; right turn; 140 miles per hour.\n\nFigure 14.- Continued.", "timestamp": "2026-07-22T04:08:28.197011+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 40, "total_pages": 104, "image_filename": "19930085842_p40.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:08:40.489946+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 16, "total_pages": 67, "image_filename": "19930085965_p16.jpg", "text": "Results of Analysis\n\nThe eddy-current power $P_e$ absorbed as heat per square centimeter of plate surface will be the product of the voltage, the current, and the power factor (fig. 6) and may be expressed as\n\n$$\nP_e = E_2 I_2 \\sin(45^\\circ + \\beta_2 - \\delta_2)\n$$\n\nor from equations (16a) and (17a)\n\n$$\nP_e = f(10^{-8}) \\frac{B_{\\max 1}}{c} \\sqrt{1 + 2 \\cos(45^\\circ + \\alpha_2) \\left( \\frac{H_{\\max 2}}{H_{\\max 1}} - 1 \\right)} \\frac{H_{\\max 2}}{\\sqrt{2} \\cdot 0.4} \\sin(45^\\circ + \\beta_2 - \\delta_2)\n$$\n\nRevision yields\n\n$$\nP_e = \\frac{1.25 \\sqrt{2} f H_{\\max 2} B_{\\max 1} 10^{-8}}{c} \\sqrt{1 + 2 \\cos(45^\\circ + \\alpha_2) \\left( \\frac{H_{\\max 2}}{H_{\\max 1}} - 1 \\right)} \\sin(45^\\circ + \\beta_2 - \\delta_2)\n$$\n\n(20)\n\nReplacement of $c$ by its parts yields\n\n$$\nP_e = \\frac{1.25 \\sqrt{3}}{\\pi} \\frac{1}{\\rho^{1/2} f^{1/2}} H_{\\max 2} B_{\\max 1} 10^{-4} \\sqrt{ \\frac{1}{\\mu} \\left[ 1 - 2 \\cos(45^\\circ + \\alpha_2) \\right] + 2 \\cos(45^\\circ + \\alpha_2) \\frac{H_{\\max 2}}{B_{\\max 1}} } \\sin(45^\\circ + \\beta_2 - \\delta_2)\n$$\n\nwatts per square centimeter  \n(20a)\n\nwhere $\\rho$ is resistivity (ohm-cm). Multiplying equation (20) by 6.45 gives the equation for the heat generated per square inch,\n\n$$\nP_e' = \\frac{11.4 f H_{\\max 2} B_{\\max 1} 10^{-8}}{c} \\sqrt{1 + 2 \\cos(45^\\circ + \\alpha_2) \\left( \\frac{H_{\\max 2}}{H_{\\max 1}} - 1 \\right)} \\sin(45^\\circ + \\beta_2 - \\delta_2), \\text{ watts per square inch}\n$$\n\n(20b)\n\nNACA RM E52G06", "timestamp": "2026-07-22T04:08:44.631284+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 12, "total_pages": 32, "image_filename": "19930085982_p12.jpg", "text": "10\nNACA RM E9E13\n\nThe design values of weight flow and pressure ratio are based\non nonviscous ideal-fluid flow. The design conditions at the inlet\nof the rotor will therefore occur at a slightly lower flow than the\ndesign value because of the boundary-layer build-up through the\ninlet bellmouth and the guide vanes. Measurements indicated that\nthe displacement thickness of the boundary layer downstream of the\nguide vanes is of the order of 0.050 inch at the tip and 0.010 inch\nat the hub. This displacement thickness will cause design angle of\nattack on the rotor blades to be theoretically obtained at a weight\nflow of 21.1 instead of 21.5 pounds per second. At this weight\nflow the over-all efficiency (extrapolated) is about 0.70. Appli-\ncation of this efficiency to the isentropic pressure ratio gives a\npredicted polytropic pressure ratio of 1.07, whereas the experimental\npressure ratio was 1.03 for the corrected design weight flow of\n21.1 pounds per second.\n\nFor a given rotor design, the pressure ratio obtained will be\na function of the relative Mach number and relative angle of attack\nat the rotor inlet. At a given rotor speed, these factors will be\ndetermined by the weight flow, the axial velocity, and the turning-\nangle distribution through the inlet guide vanes. The radial vari-\nation of relative Mach number and equivalent angle of attack at the\nrotor-blade inlet for a weight flow of 21.1 pounds per second (design\nflow) is shown in figure 5. The equivalent angle of attack was based\non an equivalent diagram having a constant axial velocity and, as\nrecommended in reference 4, utilized the mean of the inlet and outlet\naxial velocities and the true tangential components of velocity.\nFor comparison, the design values of Mach number and angle of\nattack for three-fourths design speed are also plotted in figure 5.\nGood agreement between the design and actual distributions was\nobtained. The inlet Mach number at the blade tip for a weight flow\nof 21.1 pounds per second is only slightly higher than design and\nat the compressor hub it is slightly lower than design. A compar-\nison of the experimental equivalent angle of attack with design\nvalues indicates that for the weight flow of 21.1 pounds per second,\nthe experimental angles of attack at the hub were slightly above\ndesign values whereas those at the tip were slightly below. It is\nevident that the energy addition anticipated for the given blade\ninlet Mach number and angle of attack was not obtained inasmuch as\nthe discrepancy in angle of attack is small. Figure 6 is a plot of\nmean (mass-averaged) total enthalpy addition against corrected\nweight flow. This curve shows that the design enthalpy addition\nof 90,100 foot-pounds per slug was obtained at a weight flow of\nabout 19.1 pounds per second instead of 21.1 pounds per second.", "timestamp": "2026-07-22T04:08:46.469885+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 33, "total_pages": 47, "image_filename": "19930085918_p33.jpg", "text": "32\nNACA RM A9D29\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nPressure coefficient, $P$\n-40\n-36\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\n-20\n-16\n-12\n-8\n-4\n0\n4\n\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\nChordwise station, x/c\n2\n4\n6\n8\n10\n\nSpanwise\nstation, 2y/b\n28.1%\n\n57.4%\n\n85.0%\n\n(b) $\\alpha=6.3^\\circ$\n\nNACA\n\nFigure 9.-Continued.", "timestamp": "2026-07-22T04:08:54.245156+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 58, "total_pages": 65, "image_filename": "19930085843_p58.jpg", "text": "96\n\nSemispan model\nSemispan model (Check)\nSting model\nWing-flow model\n\n$$\n\\left( \\frac{\\partial C_m}{\\partial C_L} \\right)_M\n$$\n\n0\n-.1\n-.2\n\n.6 .7 .8 .9 1.0\nMach number, M\n\nNACA\n\nFigure 17.- A comparison of $$ \\left( \\frac{\\partial C_m}{\\partial C_L} \\right)_M $$ variation with Mach number for the low-lift-coefficient range as obtained by three different test methods on models of a tailless airplane. Vertical fins on; $$ \\delta_a = 0^\\circ $$.\n\nNACA RM L9C31", "timestamp": "2026-07-22T04:08:56.294533+00:00"}
{"citation_id": "19930085927", "source_url": "https://ntrs.nasa.gov/api/citations/19930085927/downloads/19930085927.pdf", "page_number": 30, "total_pages": 30, "image_filename": "19930085927_p30.jpg", "text": "```markdown\nNACA-Langley - 8-24-49 - 275\n\n<!-- Image (126, 89, 879, 835) -->\n\nFigure 16.- Summary chart of principal hydrodynamic qualities of a flying boat having a hull length-beam ratio of 20. Gross load, 75,000 pounds; power loading, 11.5 pounds per brake horsepower; wing loading, 41.1 pounds per square foot; flap deflection, 20°.\n\nNACA RM L50C05\n\n62\n```", "timestamp": "2026-07-22T04:09:00.286055+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 29, "total_pages": 59, "image_filename": "19930085936_p29.jpg", "text": "28\nNACA RM No. E9B03\n\n<!-- Image (149, 110, 865, 854) -->\n\n(b) $\\theta = 45^\\circ$ longitudinal plane.\nFigure 5. - Continued. Pressure distributions along longitudinal planes at $0^\\circ$ yaw angle for range of angles of attack.", "timestamp": "2026-07-22T04:09:01.519003+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 7, "total_pages": 34, "image_filename": "19930086022_p7.jpg", "text": "NACA RM L9E24\n5\n\n$C_{h\\alpha}$ rate of change of aileron hinge-moment coefficient with angle of attack\n\n$P_{R\\delta}$ rate of change of aileron balance-chamber-pressure coefficient with aileron deflection\n\n$P_{R\\alpha}$ rate of change of aileron balance-chamber-pressure coefficient with angle of attack\n\n$C_{h\\delta}'$ rate of change of aileron hinge-moment coefficient in steady roll with aileron deflection\n\nMODEL AND TESTS\n\nModel\n\nThe wing was constructed of laminated mahogany to the plan form shown in figure 1. The sweep angle of the leading edge was $42.05^\\circ$, and the airfoil sections perpendicular to the 0.273 chord line were NACA $64_1$-112 sections. (The 0.273 chord line corresponds to the 0.25 chord line of the wing with unswept panels.) The aspect ratio of the wing was 4.01 and the taper ratio was 0.625. The wing tips were parabolic in plan form and elliptical in cross section. The wing was constructed with no geometric dihedral or twist.\n\nThe wing was fitted with a 0.20c' half-span outboard aileron installed on the left wing panel only. The aileron was of an internally sealed unbalanced type and had a flat-sided contour with a $11.1^\\circ$ trailing-edge angle measured perpendicular to the hinge line. The aileron hinge moments and aileron loads were measured by resistance-type electrical strain gages. The aileron seal was attached in such manner that moments and forces transmitted to the aileron were negligible. Except for cutouts to allow for the installation of strain gages, the seal extended the full span of the aileron. A total of 12 pressure orifices were installed in the balance chamber, six above and six below the seal to provide a measurement of the pressure differential across the seal. Details of the aileron are shown in figure 1.\n\nThe leading-edge flaps used in the investigation were of the round-nose extensible type with a constant chord of 3.19 inches and extended from 40 to 97.5 percent of the semispan. The deflection of the flaps was approximately $50^\\circ$ with respect to the section chord perpendicular to the 0.273 chord line.", "timestamp": "2026-07-22T04:09:01.877884+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 40, "total_pages": 92, "image_filename": "19930085930_p40.jpg", "text": "NACA RM L9G07\n39\n\n[Figure: Diagram showing a curved surface with flow lines. Labels include:\n- $M = 1.71$ with an arrow pointing to the flow.\n- $15^\\circ$ indicating an angle.\n- \"Original surface for model 4\" pointing to a dashed line on the curve.\n- Legend: Expansion (dashed line), Compression (solid line).\n- Stamps: \"CONFIDENTIAL\" and \"UNCLASSIFIED\" in red ink.\n- NACA logo at the bottom right of the diagram.]\n\n(b) Boundary-layer separation taken into consideration.\nFigure 10.- Concluded.", "timestamp": "2026-07-22T04:09:03.873022+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 90, "total_pages": 122, "image_filename": "19930082090_p90.jpg", "text": "88\nNACA TN No. 1455\n\n<!-- Image (152, 119, 802, 912) -->\n\nFigure 43.- Thermal output and isothermal frictional pressure drops of hollow-pin heat exchanger K.", "timestamp": "2026-07-22T04:09:07.071272+00:00"}
{"citation_id": "19930085917", "source_url": "https://ntrs.nasa.gov/api/citations/19930085917/downloads/19930085917.pdf", "page_number": 36, "total_pages": 38, "image_filename": "19930085917_p36.jpg", "text": "NACA RM No. L9C18 CONFIDENTIAL 35\n\nM = 0.70 M = 0.70\n\nM = 0.75 M = 0.75\n\nM = 0.81 M = 0.81\n\nNACA 64₁-012 airfoil NACA 64₁A012 airfoil\n\n(b) α = 4°.\n\nFigure 6.— Continued.\n\nCONFIDENTIAL\n\nNACA\n\nL-59836", "timestamp": "2026-07-22T04:09:09.321049+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 18, "total_pages": 43, "image_filename": "19930085958_p18.jpg", "text": "NACA RM No. L9B11\n17\n\nTABLE II\nSUMMARY OF LONGITUDINAL STABILITY CHARACTERISTICS OF\n42° SWEPTBACK WING-FUSELAGE COMBINATION WITH HORIZONTAL TAIL\n\n| Configuration | Tail height (percent b/2 above chord plane extended) | $C_m$-curve | $d\\epsilon/d\\alpha$ measured values in low lift range | $dC_m/di_t$ at $C_L=0$ |\n| :--- | :--- | :--- | :--- | :--- |\n| Flaps off low wing | Tail off | [Graph: $C_m$ vs $\\alpha$] | - - - - | - - - - |\n| | 46.6 | [Graph] | 0.32 | -0.0160 |\n| | 33.9 | [Graph] | 0.45 | -0.0160 |\n| | 21.1 | [Graph] | 0.45 | -0.0160 |\n| | -1.1 | [Graph] | 0.45 | -0.0153 |\n| 0.60b/2 drooped nose deflected 30°, split flaps, and fences. Low wing. | Tail off | [Graph] | - - - - | - - - - |\n| | 46.6 | [Graph] | 0.40 | -0.0176 |\n| | 33.9 | [Graph] | 0.43 | -0.0170 |\n| | 21.1 | [Graph] | 0.48 | -0.0166 |\n| | -1.1 | [Graph] | 0.48 | -0.0147 |\n| 0.55b/2 extensible leading edge flaps, split flaps, and fences. Low wing. | Tail off | [Graph] | - - - - | - - - - |\n| | 46.6 | [Graph] | 0.45 | -0.0173 |\n| | 33.9 | [Graph] | 0.45 | -0.0165 |\n| | 21.1 | [Graph] | 0.48 | -0.0155 |\n| | -1.1 | [Graph] | 0.45 | -0.0116 |\n| 0.70b/2 extensible leading edge flaps, split flaps, and fences. Low wing. | Tail off | [Graph] | - - - - | - - - - |\n| | 46.6 | [Graph] | 0.36 | -0.0170 |\n| | 33.9 | [Graph] | 0.48 | -0.0168 |\n| | 21.1 | [Graph] | 0.48 | -0.0160 |\n| | -1.1 | [Graph] | 0.50 | -0.0150 |\n\nNACA", "timestamp": "2026-07-22T04:09:10.574535+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 11, "total_pages": 54, "image_filename": "19930086015_p11.jpg", "text": "```markdown\n10 CONFIDENTIAL NACA RM A9E24\n\nIf the pressure distribution over either the upper or lower surface of the wing is required independently, however, the influences of stream curvature in producing effective camber and twist must be minimized by mounting the model with the span vertical. There is also the question of correcting for the pressure variation in the stream so that the pressures over the model may be referred to the average ambient pressure of the stream. This may be done by a simple superposition process as in reference 4. The correction can be reasoned as a valid first-order approximation if the pressure disturbances (expansion or compression waves of infinitesimal strength) are small and are not reflected by the wing surface. If the flow in the nozzle is two-dimensional, reflection of the pressure disturbances will not occur if the plane of the model wing is placed so as to intersect at right angles the planes along which the weak pressure waves are propagated. In the Ames 6- by 6-foot supersonic wind-tunnel nozzle, the plane of the model wing must be placed vertically to insure the validity of the correction.\n\nCONCLUDING REMARKS\n\nThe surveys of the air stream in the test section of the asymmetric adjustable nozzle of the Ames 6- by 6-foot supersonic wind tunnel show that the flow is nearly uniform at a nominal Mach number of 1.4. As the Mach number is increased or decreased from a value of 1.4, however, vertical pressure gradients of significant magnitude are found. Smaller axial gradients also exist. The transverse gradients are of negligible magnitudes which indicates that the flow is essentially two-dimensional.\n\nThe existence of large vertical pressure gradients implies an appreciable variation in stream angle with axial position. Stream-angle measurements confirm this.\n\nThe results of tests of one swept-wing model indicated that for this model, at least, the effects of the nonuniformity of the stream on certain model characteristics may be minimized by testing with the plane of the model wing parallel to the two-dimensional-flow planes. For other model characteristics which are combined functions of the angle of pitch and the angle of yaw, this method will not be effective. In such cases, appropriate corrections need to be applied. At some Mach numbers, the magnitude and uncertainty of these corrections may be such as to preclude certain tests. Research devoted to the refinement of nozzle design techniques is now proceeding with a view toward improving the flow in the wind tunnel at these Mach numbers.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:09:13.490504+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 57, "total_pages": 118, "image_filename": "19930085838_p57.jpg", "text": "```markdown\nNACA RM No. L9B23\n55\n\n<!-- Image (159, 119, 877, 909) -->\n\n(1) $\\delta_F = 40^\\circ$.\nFigure 7.- Continued.\n```", "timestamp": "2026-07-22T04:09:13.893541+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 12, "total_pages": 36, "image_filename": "19930086003_p12.jpg", "text": "10\nCONFIDENTIAL\nNACA RM L9I08\n\nTABLE I.- FUSELAGE ORDINATES\n\nBasic fineness ratio 12; actual fineness ratio 10\nachieved by cutting off the rear one-sixth of\nthe body; 5/4 located at 1/2\n\n[Figure: Diagram of a fuselage shape with dimensions labeled: l = 14.14, 5/6 l, 1/2, x, r, D(Max)]\n\nOrdinates\n\n| x/l | r/l | x/l | r/l |\n| :--- | :--- | :--- | :--- |\n| 0 | 0 | 0 | 0 |\n| .005 | .00231 | .4500 | .04143 |\n| .0075 | .00298 | .5000 | .04167 |\n| .0125 | .00428 | .5500 | .04130 |\n| .0250 | .00722 | .6000 | .04024 |\n| .0500 | .01205 | .6500 | .03842 |\n| .0750 | .01613 | .7000 | .03562 |\n| .1000 | .01971 | .7500 | .03128 |\n| .1500 | .02593 | .8000 | .02526 |\n| .2000 | .03090 | .8338 | .02000 |\n| .2500 | .03465 | .8500 | .01852 |\n| .3000 | .03741 | .9000 | .01125 |\n| .3500 | .03933 | .9500 | .00439 |\n| .4000 | .04063 | 1.0000 | 0 |\n\nL. E. radius = 0.0005l\n\nNACA\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:09:17.926020+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 26, "total_pages": 50, "image_filename": "19930085952_p26.jpg", "text": "NACA RM L9C24\n25\n\n<!-- Image (208, 109, 838, 874) -->\n\nFigure 3.- Blade-form curves for the model propeller.", "timestamp": "2026-07-22T04:09:25.412861+00:00"}
{"citation_id": "19930085842", "source_url": "https://ntrs.nasa.gov/api/citations/19930085842/downloads/19930085842.pdf", "page_number": 41, "total_pages": 104, "image_filename": "19930085842_p41.jpg", "text": "NACA RM L9029\n37\n\n[Figure: A graph with two plots. The top plot shows Propeller advance-diameter ratio, V/nD on the y-axis (0 to 10) against Torque coefficient, Qc on the x-axis (0 to 24). The curve is labeled V/nD. A horizontal line is labeled 1200 bhp at 1085 rpm. The bottom plot shows Lift coefficient, CL on the y-axis (0 to 12) against Torque coefficient, Qc on the x-axis (0 to 24). The line is labeled CL. The bottom right corner contains the text: NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS.]\n\nFigure 11.- Variation of $C_L$ and $V/nD$ with $Q_c$ of the airplane for full-power operation at sea level.", "timestamp": "2026-07-22T04:09:31.781757+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 56, "total_pages": 82, "image_filename": "19930085551_p56.jpg", "text": "NACA RM No. L8K30\n55\n\n[Figure: A series of five stacked time-series graphs plotting flight data. The x-axis for all graphs is \"Time, sec\" ranging from 0 to 40. The top graph shows \"Yawing velocity\" and \"Pitching velocity\". The second graph shows \"Control force, lb\". The third graph shows \"Control position, deg\". The fourth graph shows \"Altitude, ft\". The bottom graph shows \"Indicated airspeed, mph\". A NACA logo is present in the bottom right corner of the bottom graph.]\n\n(f) Approach condition; flaps $20^\\circ$; gear down; power for level flight; altitude controlled; right turn; 140 miles per hour.\n\nFigure 14.— Continued.", "timestamp": "2026-07-22T04:09:33.123685+00:00"}
{"citation_id": "19930085965", "source_url": "https://ntrs.nasa.gov/api/citations/19930085965/downloads/19930085965.pdf", "page_number": 17, "total_pages": 67, "image_filename": "19930085965_p17.jpg", "text": "16\nNACA RM E9E06\n\nOperation at Constant Permeability\n\nThe equations that have been derived are not limited in use to conditions in which the outer plate surface is saturated, but are also applicable in the region of the approximated magnetization curve where the permeability $\\mu$ is constant. At the saturation point, where $H_{max \\ 2}$ is equal to $H_{max \\ 1}$, the equation for the total depth of penetration (equation (15c)) reduces to equation (7a). The depth of penetration remains constant at the value given by equation (7a), even if the impressed field is reduced below $H_{max \\ 1}$. The power-factor angle $\\theta$ remains $45^\\circ$ for field intensities equal to or less than $H_{max \\ 1}$, and $\\beta_2$ and $\\delta_2$ vanish. At the saturation point, the radical in the voltage equation (16a), the total-flux equation (10b), and the power equation (20b) reduces to a value of 1. These equations are still applicable for impressed field intensities less than $H_{max \\ 1}$ with the substitution of the impressed field intensity $H_{max \\ s}$ for $H_{max \\ 2}$ and the value equivalent to the impressed field $R_{max \\ s}$ for $R_{max \\ 1}$ after the radical has been replaced by 1. The total-flux equation (10b) reduces to equation (4a). The current formula (17a) applies if the given impressed field intensity $H_{max \\ s}$ is used for $H_{max \\ 2}$.\n\nEXPERIMENTAL INVESTIGATION\n\nPrior to any attempt at full-scale application of eddy-current heating of the inlet guide vanes, preliminary investigations were conducted from which information was obtained as to the number of field ampere turns required to produce specified amounts of heat.\n\nThe schematic layout of the equipment in which sample blades were investigated is shown in figure 10. For simplicity, instead of using a chopper to vary the flux, a similar effect was accomplished by varying the magnetomotive force or ampere turns by connecting the coil shown in figure 10 to a 500-watt, 6100-cycle-per-second, audio-frequency power generator. In order to match the load to the generator properly, the turns on the coil were varied from 316 to 210, in addition to resonating this load with series capacitors. The current in the coil was measured with radio-frequency ammeters.\n\nThe heat generated was determined by measuring the temperature rise and the rate of water flowing through the jacket. Heat exchange to the laminated core was minimized by keeping the core", "timestamp": "2026-07-22T04:09:42.835245+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 59, "total_pages": 65, "image_filename": "19930085843_p59.jpg", "text": "NACA RM L59J31\n\nSemispan model\nSting model\nWing-flow model\n\n$$\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_{M=1}$$\n\n[Graph: Plot of $\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_{M=1}$ vs. Mach number, M, with three curves representing Semispan model, Sting model, and Wing-flow model. X-axis ranges from 0.4 to 1.0. Y-axis ranges from -2 to 0. NACA logo in lower right corner of plot area.]\n\nMach number, M\n\nFigure 18.— A comparison of $\\left(\\frac{\\partial C_m}{\\partial C_L}\\right)_M$ variation with Mach number for the low-lift-coefficient range as obtained by three different test methods on models of a tailless airplane. Vertical fins off; $\\delta_a = 0^\\circ$.\n\n57", "timestamp": "2026-07-22T04:09:44.369130+00:00"}
{"citation_id": "19930085588", "source_url": "https://ntrs.nasa.gov/api/citations/19930085588/downloads/19930085588.pdf", "page_number": 44, "total_pages": 51, "image_filename": "19930085588_p44.jpg", "text": "NACA RM No. L51D08\n\n| Moment coefficient, $c_m,c/4$ | Section lift coefficient, $c_l$ | Section drag coefficient, $c_d$ | Moment coefficient, $c_m,a.c.$ |\n| :--- | :--- | :--- | :--- |\n| -2.8 | | .032 | |\n| -2.4 | | .028 | |\n| -2.0 | | .024 | |\n| -1.6 | | .020 | |\n| -1.2 | | .016 | |\n| -.8 | | .012 | |\n| -.4 | | .008 | |\n| 0 | | .004 | |\n| +.4 | | .000 | |\n| +.8 | | | |\n| +1.2 | | | |\n| +1.6 | | | |\n| +2.0 | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n| | | | |\n|", "timestamp": "2026-07-22T04:09:44.823657+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 30, "total_pages": 59, "image_filename": "19930085936_p30.jpg", "text": "NACA RM No. E9B03\n29\n\nPressure coefficient, $C_p$\nAngle of attack (deg)\n$\\circ$ -15\n$\\square$ -6\n$\\diamond$ 0\n$\\triangle$ 12\n$\\nabla$ 24\n\nDistance from tip, x/L\n\n(c) $\\theta = 180^\\circ$ longitudinal plane.\n\nFigure 5. - Continued. Pressure distributions along longitudinal planes at $0^\\circ$ yaw angle for range of angles of attack.\n\nNACA", "timestamp": "2026-07-22T04:09:47.378437+00:00"}
{"citation_id": "19930085930", "source_url": "https://ntrs.nasa.gov/api/citations/19930085930/downloads/19930085930.pdf", "page_number": 41, "total_pages": 92, "image_filename": "19930085930_p41.jpg", "text": "40\n\nUNCLASSIFIED\nCONFIDENTIAL\n\n[Figure: Diagram showing two curved blade shapes, one above the other, with an arrow pointing to the left side of the lower shape. The NACA logo is visible near the bottom right of the diagram.]\n\nUNCLASSIFIED\nCONFIDENTIAL\n\n(a) Blade corresponding to figure 10(a). Stagger angle, $47^\\circ$.\n\nFigure 11.- Possible blade shapes for model 4.\n\nNACA RM L59G07", "timestamp": "2026-07-22T04:09:50.379724+00:00"}
{"citation_id": "19930086060", "source_url": "https://ntrs.nasa.gov/api/citations/19930086060/downloads/19930086060.pdf", "page_number": 1, "total_pages": 20, "image_filename": "19930086060_p1.jpg", "text": "```markdown\n386\nCopy\nRM L9F02\n\nFILE COPY\nNO 3\n\nNACA RM L9F02\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nFLIGHT INVESTIGATION AT HIGH-SUBSONIC, TRANSONIC,\nAND SUPERSONIC SPEEDS TO DETERMINE ZERO-LIFT DRAG OF\nBODIES OF REVOLUTION HAVING FINENESS RATIO OF 6.04\nAND VARYING POSITIONS OF MAXIMUM DIAMETER\n\nBy\n\nEllis R. Katz\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\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 50:31 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: CROWLEY CHANGE #2236\nDATE 1-8-54 T.C.F.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nAugust 31, 1949\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T04:09:53.067102+00:00"}
{"citation_id": "19930085982", "source_url": "https://ntrs.nasa.gov/api/citations/19930085982/downloads/19930085982.pdf", "page_number": 13, "total_pages": 32, "image_filename": "19930085982_p13.jpg", "text": "NACA RM E9E13\n\nApproach to equilibrium. - There are two common methods of determining the radial distribution of axial velocity at the inlet to a stator-blade row. One is the assumption that the axial velocity distribution at the stator inlet is the same as that at the rotor inlet. This condition is for zero radial flow for an incompressible process. The absolute magnitude of the axial velocities are then determined from continuity. By this assumption, radial accelerations are admitted, but the elapsed time for a particle of air to pass from the rotor inlet to the stator inlet is assumed to be sufficiently small to render the radial flow negligible. The second method is to assume that simple-radial equilibrium of static pressure and centrifugal forces exists. In this case the velocities at the stator inlet are determined from the simple-radial-equilibrium equation\n\n$$\n\\frac{1}{\\rho} \\frac{dp}{dr} = \\frac{V_{\\theta}^2}{r}\n$$\n\nin conjunction with continuity and the energy equation. Radial accelerations within the blade row are assumed to be sufficiently large to establish the static pressure for simple equilibrium at the stator inlet. The complexity of the complete Euler equation for radial equilibrium makes it expedient to use one of these two assumptions to obtain a simplified design theory. The actual flow distribution immediately behind the rotor will, however, probably fall somewhere between that assuming no change in axial velocity through the rotor row and that for simple equilibrium. Simple equilibrium will be more closely approached as the distance downstream of the rotor is increased.\n\nAs was shown in the section Compressor Design, this compressor was designed using the assumption of constant axial velocity within the blade row.\n\nIn order to determine to what extent the actual flow followed the design assumption of constant axial velocity over the blade row, curves are plotted in figure 7 comparing the measured axial-velocity distribution (in terms of velocity ratio $V_z/\\sqrt{V_{z,m}}$) entering the rotor at three-fourths design speed and a flow of 18.2 pounds per second with that obtained at the measuring station downstream of the rotor. The axial-velocity distribution that would be expected at the rotor outlet if the flow were in simple-radial equilibrium using the measured tangential velocities is also plotted. It is apparent that some radial shift in the flow occurred so as to change the axial-velocity distribution over the blade row. This", "timestamp": "2026-07-22T04:09:56.765573+00:00"}
{"citation_id": "19930085529", "source_url": "https://ntrs.nasa.gov/api/citations/19930085529/downloads/19930085529.pdf", "page_number": 73, "total_pages": 85, "image_filename": "19930085529_p73.jpg", "text": "```markdown\n72\nNACA RM No. L8A30a\n\nTABLE 66\n$$\n\\left[ \\Lambda = -30^\\circ, \\delta_{a_u} = 5.0^\\circ, \\alpha = 4^\\circ \\right]\n$$\n\nCONFIDENTIAL\n\n| UPPER SURFACE | | | | | | | | LOWER SURFACE | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| **Tube** | **Per-cent chord** | **Mach Number** | | | | | | **Tube** | **Per-cent chord** | **Mach Number** | | | | | |\n| | | **0.60** | **0.80** | **0.89** | **0.925** | **0.96** | | | | **0.60** | **0.80** | **0.89** | **0.925** | **0.96** |\n| A 1 | 2.0 | -- | -- | -- | -- | -- | -- | 86 | 3.0 | -- | -- | -- | -- | -- | -- |\n| 2 | 5.0 | -- | -- | -- | -- | -- | -- | 87 | 10.0 | -- | -- | -- | -- | -- | -- |\n| 3 | 15.0 | -- | -- | -- | -- | -- | -- | 88 | 25.0 | -- | -- | -- | -- | -- | -- |\n| 4 | 27.5 | -- | -- | -- | -- | -- | -- | 89 | 41.0 | -- | -- | -- | -- | -- | -- |\n| 5 | 40.0 | -- | -- | -- | -- | -- | -- | 90 | 52.5 | -0.064 | -0.098 | -0.102 | -0.142 | -0.193 | -0.382 |\n| 6 | 50.0 | -0.299 | -0.464 | -0.544 | -0.544 | -0.369 | -0.705 | 91 | 62.5 | -.023 | -.043 | -.077 | -.103 | -.110 | -.309 |\n| 7 | 60.0 | -.182 | -.395 | -.490 | -.590 | -.290 | -.636 | 92 | 72.5 | -.020 | -.095 | -.086 | -.078 | -.086 | -.318 |\n| 8 | 67.5 | -.130 | -.284 | -.452 | -.506 | -.307 | -.457 | 93 | 84.0 | -- | -- | -- | -- | -- | -- |\n| 9 | 77.5 | -- | -- | -- | -- | -- | -- | 94 | 94.0 | -- | -- | -- | -- | -- | -- |\n| 10 | 87.5 | -- | -- | -- | -- | -- | -- | | | | | | | | |\n| 11 | 96.0 | -- | -- | -- | -- | -- | -- | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| B12 | 2.0 | -1.184 | -1.233 | -1.079 | -.992 | -.765 | -.626 | 95 | 3.0 | .597 | .568 | .549 | .490 | .465 | .432 |\n| 13 | 5.0 | -1.165 | -1.177 | -1.009 | -.968 | -.804 | -.747 | 96 | 10.0 | .271 | .258 | .235 | .215 | .180 | .189 |\n| 14 | 15.0 | -.831 | -.840 | -.682 | -.702 | -.711 | -.816 | 97 | 25.0 | .031 | .000 | -.030 | -.057 | -.107 | -.119 |\n| 15 | 27.5 | -.618 | -.599 | -.547 | -.563 | -.894 | -.865 | 98 | 41.0 | -.060 | -.101 | -.138 | -.190 | -.271 | -.284 |\n| 16 | 40.0 | -.456 | -.289 | -.216 | -.299 | -.774 | -.886 | 99 | 52.5 | -.067 | -.106 | -.143 | -.188 | -.288 | -.353 |\n| 17 | 50.0 | -.376 | -.219 | -.197 | -.262 | -.638 | -.826 | 100 | 62.5 | -.050 | -.062 | -.096 | -.127 | -.203 | -.310 |\n| 18 | 59.0 | -.286 | -.105 | -.100 | -.168 | -.576 | -.852 | 101 | 72.5 | -.020 | -.031 | -.064 | -.084 | -.144 | -.174 |\n| 19 | 67.5 | -.193 | -.293 | -.308 | -.374 | -.524 | -.795 | 102 | 86.3 | .084 | .073 | .036 | .019 | .000 | -.127 |\n| 20 | 77.5 | -.096 | -.182 | -.247 | -.288 | -.468 | -.725 | 103 | 94.4 | -.130 | -.115 | -.058 | -.031 | .016 | -.051 |\n| 21 | 88.0 | -.010 | -.051 | -.156 | -.230 | -.366 | -.521 | | | | | | | | |\n| 22 | 95.5 | -- | -- | -- | -- | -- | -- | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| C23 | 2.0 | -1.453 | -1.072 | -.831 | -.643 | -.493 | -.378 | 104 | 3.0 | .555 | .542 | .525 | .495 | .471 | .468 |\n| 24 | 5.0 | -.931 | -1.147 | -.949 | -.780 | -.647 | -.536 | 105 | 10.0 | -.270 | .271 | .258 | .236 | .219 | .205 |\n| 25 | 15.0 | -.700 | -1.102 | -.921 | -.804 | -.689 | -.585 | 106 | 25.0 | .043 | .051 | .032 | .007 | -.014 | -.006 |\n| 26 | 27.8 | -.606 | -.592 | -.621 | -.804 | -.763 | -.664 | 107 | 41.0 | -.033 | -.066 | -.082 | -.116 | -.168 | -.167 |\n| 27 | 40.0 | -.531 | -.518 | -.577 | -.842 | -.820 | -.701 | 108 | 52.5 | -.068 | -.088 | -.117 | -.164 | -.240 | -.271 |\n| 28 | 50.0 | -.436 | -.404 | -.596 | -.737 | -.777 | -.783 | 109 | 62.5 | -.032 | -.054 | -.079 | -.110 | -.196 | -.238 |\n| 29 | 59.0 | -.350 | -.346 | -.508 | -.646 | -.777 | -.876 | 110 | 72.5 | -.017 | -.036 | -.060 | -.085 | -.139 | -.143 |\n| 30 | 67.5 | -.232 | -.256 | -.277 | -.545 | -.731 | -.791 | 111 | 85.1 | .071 | .093 | .073 | .061 | .009 | -.058 |\n| 31 | 77.5 | -.127 | -.151 | -.130 | -.287 | -.570 | -.840 | 112 | 94.4 | .121 | .136 | .122 | .115 | .046 | -.018 |\n| 32 | 88.0 | -.013 | -.015 | -.041 | -.089 | -.247 | -.320 | | | | | | | | |\n| 33 | 95.5 | -- | -- | -- | -- | -- | -- | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| D34 | 2.0 | -1.217 | -1.000 | -.789 | -.609 | -.474 | -.370 | 113 | 3.0 | .513 | .508 | .491 | .461 | .442 | .429 |\n| 35 | 15.0 | -.659 | -1.025 | -.893 | -.760 | -.646 | -.547 | 114 | 10.0 | .263 | .270 | .262 | .242 | .229 | .233 |\n| 36 | 27.5 | -.559 | -.552 | -.608 | -.777 | -.765 | -.665 | 115 | 25.0 | .043 | .051 | .032 | .007 | -.014 | -.006 |\n| 37 | 40.0 | -.528 | -.508 | -.564 | -.908 | -.818 | -.733 | 116 | 41.0 | -.029 | -.045 | -.065 | -.101 | -.136 | -.136 |\n| 38 | 50.0 | -.453 | -.418 | -.564 | -.842 | -.820 | -.811 | 117 | 52.5 | -.068 | -.088 | -.117 | -.164 | -.240 | -.271 |\n| 39 | 59.0 | -.350 | -.430 | -.713 | -.847 | -.837 | -.802 | 118 | 62.5 | -.032 | -.033 | -.057 | -.104 | -.196 | -.173 |\n| 40 | 67.5 | -- | -- | -- | -- | -- | -- | 119 | 72.5 | -.023 | -.031 | -.051 | -.071 | -.013 | -.144 |\n| 41 | 77.5 | -.140 | -.136 | -.120 | -.263 | -.525 | -.616 | 120 | 87.4 | .077 | .148 | .109 | .067 | -.020 | -.020 |\n| 42 | 87.5 | -.033 | -.087 | -.082 | -.213 | -.391 | -.301 | 121 | 94.2 | .116 | .111 | .107 | .068 | -.012 | .017 |\n| 43 | 94.2 | -.021 | .038 | .056 | .010 | -.055 | -.086 | | | | | | | | |\n| | | | | | | | | | | | | | | | |\n| E44 | 2.0 | -1.571 | -1.110 | -.879 | -.699 | -.537 | | 122 | 3.0 | .562 | .559 | .547 | .528 | .508 | |\n| 45 | 5.0 | -1.074 | -1.107 | -.975 | -.820 | -.679 | | 123 | 10.0 | .294 | .288 | .281 | .269 | .277 | |\n| 46 | 15.0 | -.682 | -1.103 | -.951 | -.820 | -.698 | | 124 | 25.0 | .090 | .114 | .089 | .075 | .055 | |\n| 47 | 27.5 | -.556 | -.552 | -.608 | -.777 | -.765 | | 125 | 41.0 | .001 | -.013 | -.029 | -.058 | -.010 | |\n| 48 | 40.0 | -.538 | -.480 | -.533 | -.878 | -.720 | | 126 | 52.5 | -.001 | -.006 | -.012 | -.023 | -.063 | |\n| 49 | 50.0 | -.453 | -.418 | -.564 | -.842 | -.816 | | 127 | 62.5 | .001 | -.004 | -.012 | -.023 | -.063 | |\n| 50 | 59.0 | -.362 | -.399 | -.364 | -.631 | -.669 | | 128 | 72.5 | .031 | .042 | .007 | -.005 | -.056 | |\n| 51 | 67.5 | -.260 | -.278 | -.281 | -.478 | -.550 | | 129 | 85.1 | .051 | .071 | .051 | .031 | -.014 | |\n| 52 | 77.5 | -.137 | -.136 | -.116 | -.388 | -.480 | | 130 | 86.3 | .061 | .054 | .097 | .054 | -.014 | |\n| 53 | 88.0 | -.013 | -.044 | -.043 | -.1", "timestamp": "2026-07-22T04:09:57.707256+00:00"}
{"citation_id": "19930082090", "source_url": "https://ntrs.nasa.gov/api/citations/19930082090/downloads/19930082090.pdf", "page_number": 91, "total_pages": 122, "image_filename": "19930082090_p91.jpg", "text": "NACA TN No. 1455\n89\n\n50.5\"\nExhaust gas\n15.7\"\nVentilating air\nB\nVentilating air\nB\n6.4\"\nExhaust gas\n8.4\"\nSection B-B\nUC-1 air shroud\n15.8\"\n6.5\"\n25.8\"\n24 rows of pins\n16 pins per row\n1.0\"\n0.75\"\n0.32\" diameter\nNACA\n\nFigure 44.- Schematic diagram of heat exchanger L and air shroud. Weight of heat exchanger, 25 pounds.\n\n| Minimum cross-sectional area, sq ft | Air side | Gas side |\n| :--- | :--- | :--- |\n| | 0.184 | 0.191 |", "timestamp": "2026-07-22T04:10:03.517234+00:00"}
{"citation_id": "19930086015", "source_url": "https://ntrs.nasa.gov/api/citations/19930086015/downloads/19930086015.pdf", "page_number": 12, "total_pages": 54, "image_filename": "19930086015_p12.jpg", "text": "NACA RM A9E24 CONFIDENTIAL 11\n\nREFERENCES\n\n1. Allen, H. Julian: The Asymmetric Adjustable Supersonic Nozzle for Wind-Tunnel Application. NACA RM A8E17, 1948.\n\n2. Stevens, Victor I., and Boyd, John W.: A Comparison of Theoretical and Experimental Loading on a $63^\\circ$ Swept-Back Wing at Supersonic Speeds. NACA RM A9C16, 1949.\n\n3. Chapman, Dean R., and Perkins, Edward W.: Experimental Investigation of the Effects of Viscosity on the Drag of Bodies of Revolution at a Mach Number of 1.5. NACA RM A7A31a, 1947.\n\n4. Frick, Charles W., and Boyd, John W.: Investigation at Supersonic Speed (M = 1.53) of the Pressure Distribution over a $63^\\circ$ Swept Airfoil of Biconvex Section at Zero Lift. NACA RM A8C22, 1948.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T04:10:05.749228+00:00"}
{"citation_id": "19930085957", "source_url": "https://ntrs.nasa.gov/api/citations/19930085957/downloads/19930085957.pdf", "page_number": 24, "total_pages": 27, "image_filename": "19930085957_p24.jpg", "text": "```markdown\nNACA RM L9E02\n\n1.6\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n$C_L$\n\n-4 0 4 8 12 16 20 24 28\n$\\alpha$ deg\n\n.08 .12 .16 .20 .24 .28 .32 .36\n$C_D$\n\n0 -.04 -.08 -.12 -.16\n$C_m$\n\n[Figure: Graph showing $C_L$ vs $\\alpha$, $C_L$ vs $C_D$, and $C_L$ vs $C_m$. Data points are marked with squares and triangles. Curves are labeled \"Smooth\" and \"Rough\". A NACA logo is present in the bottom right corner of the graph area.]\n\n(b) Half-span split flaps on; $\\delta_F = 60^\\circ$; $R = 3.0 \\times 10^6$.\n\nFigure 11.- Concluded.\n\n23\n```", "timestamp": "2026-07-22T04:10:05.806461+00:00"}
{"citation_id": "19930085958", "source_url": "https://ntrs.nasa.gov/api/citations/19930085958/downloads/19930085958.pdf", "page_number": 19, "total_pages": 43, "image_filename": "19930085958_p19.jpg", "text": "```markdown\n18\nNACA RM No. L9B11\n\n<!-- Image (52, 110, 888, 936) -->\n\nFUSELAGE ORDINATES\n\n| Distance behind fuselage nose | Fuselage diameter | Distance behind fuselage nose | Fuselage diameter |\n| :--- | :--- | :--- | :--- |\n| 0 | 0.20 | 112.00 | 16.80 |\n| 18.00 | 9.84 | 122.00 | 16.32 |\n| 22.05 | 11.80 | 132.00 | 14.90 |\n| 27.39 | 13.80 | 142.00 | 12.52 |\n| 34.56 | 15.60 | 151.20 | 9.46 |\n| 42.35 | 16.60 | 162.00 | 4.78 |\n| 48.00 | 16.80 | 170.95 | 0 |\n\nNACA\n\nFigure 1.- Geometry of wing, fuselage, and horizontal tail.\n```", "timestamp": "2026-07-22T04:10:08.632747+00:00"}
{"citation_id": "19930086022", "source_url": "https://ntrs.nasa.gov/api/citations/19930086022/downloads/19930086022.pdf", "page_number": 8, "total_pages": 34, "image_filename": "19930086022_p8.jpg", "text": "6\nNACA RM L9E24\n\nSplit flaps were used as the trailing-edge high-lift device in this investigation and had a chord of 20 percent of the wing chord measured perpendicular to the 0.273 chord line (which corresponds to 18.4 percent of the wing chord measured parallel to the plane of symmetry) and extended from the plane of symmetry to 50 percent of the semispan. The flaps were constructed of $\\frac{1}{16}$-inch sheet steel and were attached to the wing at an angle of 60°. The angle between the flap chord line and the lower surface of the wing thereby constitutes the angle of flap deflection and is measured perpendicular to the 0.273 chord line.\n\nThe upper-surface fences were located at 50 percent of the wing semispan and were constructed of $\\frac{1}{16}$-inch sheet steel cut to fit the upper surface of the wing. The fences extended from 5 percent of the local chord to the wing trailing edge. The height of the fences was arbitrarily set at 60 percent of the maximum thickness of the local airfoil section parallel to the plane of symmetry.\n\nDetails of the high-lift and stall-control devices are shown in figure 2.\n\nTests\n\nThe tests were conducted in the Langley 19-foot pressure tunnel with the model mounted in the tunnel as shown in figure 3. The air in the tunnel was compressed to approximately 0.0055 slugs per cubic foot enabling the tests to be made at a Reynolds number of 6,800,000 and a Mach number of 0.16.\n\nThe aileron lateral control characteristics and the wing lift, drag, and pitching-moment characteristics were determined for the plain wing and the wing equipped with high-lift and stall-control devices for an angle-of-attack range from -4° through the stall and for an aileron-deflection range from -25° to 25°.\n\nREDUCTION OF DATA\n\nAll data have been reduced to nondimensional coefficient form. Corrections for support tare and interference have been applied to all force and moment data. Jet-boundary corrections determined by means of reference 4 and air-flow-misalignment corrections have been applied to the angle of attack and drag coefficient. In addition, a", "timestamp": "2026-07-22T04:10:08.708942+00:00"}
{"citation_id": "19930086003", "source_url": "https://ntrs.nasa.gov/api/citations/19930086003/downloads/19930086003.pdf", "page_number": 13, "total_pages": 36, "image_filename": "19930086003_p13.jpg", "text": "CONFIDENTIAL\n\nNACA RM L9J08\n\n1.299\n45°\n1.768 ± ε\n5.196\n0.25 Chord line\n2.382\n90°\nReference centerline\nClearance 1/16\n2.165\nBump surface\nCenterline of balance normal to bump surface\n7.07\n11.8\n1.18 Maximum diameter\n0.56\nWing-alone end plate\n2.50\nWing-fuselage end plate\n\nTabulated Wing Data\nArea (Twice semispan) 0.125 sq ft\nMean aerodynamic chord 0.1473 ft\nAspect ratio 6\nTaper ratio 0.6\nIncidence 0.0°\nDihedral 0.0°\nAirfoil section parallel to free stream NACA 65A006\n\nScale, inches\n0 1 2\n\nNACA\n\nFigure 1.— General arrangement of a model with 45° sweptback wing, aspect ratio 6, taper ratio 0.6, and NACA 65A006 airfoil section.\n\nCONFIDENTIAL\n\n11", "timestamp": "2026-07-22T04:10:09.278543+00:00"}
{"citation_id": "19930085918", "source_url": "https://ntrs.nasa.gov/api/citations/19930085918/downloads/19930085918.pdf", "page_number": 34, "total_pages": 47, "image_filename": "19930085918_p34.jpg", "text": "NACA RM A9D29\n33\n\nPressure coefficient, P\n-76\n-72\n-68\n-64\n-60\n-56\n-52\n-48\n-44\n-40\n-36\n-32\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n4\n8\n\n-28\n-24\n-20\n-16\n-12\n-8\n-4\n0\n-16\n-12\n-8\n-4\n0\n4\n8\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n28.1%\n57.4%\n85.0%\n\nChordwise station, x/c\n0\n4\n6\n8\n10\n\n(c) $\\alpha=12.5^\\circ$\n\nFigure 9.—Continued.", "timestamp": "2026-07-22T04:10:14.906515+00:00"}
{"citation_id": "19930085952", "source_url": "https://ntrs.nasa.gov/api/citations/19930085952/downloads/19930085952.pdf", "page_number": 27, "total_pages": 50, "image_filename": "19930085952_p27.jpg", "text": "26\nNACA RM L9C24\n\n$$Q_c$$\n.04\n.03\n.02\n.01\n0\n\n$$\\beta, \\text{deg}$$\n20\n14.5\n13\n\n1200 bhp at 1085 rpm\n\n$$V/nD$$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n$$\\circ$$ Articulated propeller\n$$\\square$$ Rigid propeller\n\n$$\\beta, \\text{deg}$$\n20\n14.5\n13\n\n0 .4 .8 1.2 1.6 2.0 2.4\n$$C_L$$\n\nNACA\n\nFigure 4.- Variation of $$Q_c$$ and $$V/nD$$ with $$C_L$$ for simulated full-power operation at sea level.", "timestamp": "2026-07-22T04:10:20.791279+00:00"}
{"citation_id": "19930085551", "source_url": "https://ntrs.nasa.gov/api/citations/19930085551/downloads/19930085551.pdf", "page_number": 57, "total_pages": 82, "image_filename": "19930085551_p57.jpg", "text": "56\nNACA RM No. L8K30\n\nYawing velocity,\nLeft rad/sec\nRight rad/sec\nPitching velocity,\nUp rad/sec\nDown rad/sec\nPitch\nYaw\n\nControl forces,\nPush lb\nPull lb\nElevator\nAileron\nRudder\n\nControl position,\nLeft deg\nRight deg\nDown deg\nUp deg\nElevator\nAileron\nRudder\n\nAltitude, ft\n\nIndicated airspeed, mph\n\nTime, sec\n\n(g) Approach condition; flaps 20°; gear down; power for level flight;\nno attempt to control altitude; left turn; 140 miles per hour.\n\nFigure 14.- Continued.", "timestamp": "2026-07-22T04:10:23.782549+00:00"}
{"citation_id": "19930085843", "source_url": "https://ntrs.nasa.gov/api/citations/19930085843/downloads/19930085843.pdf", "page_number": 60, "total_pages": 65, "image_filename": "19930085843_p60.jpg", "text": "58\n\nPitching-moment coefficient, $C_{m_{\\delta=0}}$\n\nSemispan model\nSting model\nWing-flow model\nWing-flow model (check)\n\n.04\n0\n-.04\n\n.4 .5 .6 .7 .8 .9 1.0 1.1\n\nMach number, M\n\nNACA\n\n(a) Vertical fins on.\n\nFigure 19.- A comparison of pitching-moment coefficient through the Mach number range at zero lift coefficient as obtained by different test methods on models of a tailless airplane. $\\delta_a = 0^\\circ$.\n\nNACA RM L9C31", "timestamp": "2026-07-22T04:10:31.402960+00:00"}
{"citation_id": "19930085838", "source_url": "https://ntrs.nasa.gov/api/citations/19930085838/downloads/19930085838.pdf", "page_number": 58, "total_pages": 118, "image_filename": "19930085838_p58.jpg", "text": "```markdown\n56\n\nAileron section hinge-moment coefficient, $c_{h_a}$\n\n<!-- Image (57, 79, 896, 792) -->\n\n(m) $\\delta_f = 40^\\circ$\nFigure 7.-- Concluded.\n\nNACA EW No. 10923\n```", "timestamp": "2026-07-22T04:10:32.924555+00:00"}
{"citation_id": "19930085936", "source_url": "https://ntrs.nasa.gov/api/citations/19930085936/downloads/19930085936.pdf", "page_number": 31, "total_pages": 59, "image_filename": "19930085936_p31.jpg", "text": "30\nNACA RM No. E9B03\n\n<!-- Image (156, 109, 863, 859) -->\n\n(d) $\\theta = 225^\\circ$ longitudinal plane.\nFigure 5. - Continued. Pressure distributions along longitudinal planes at $0^\\circ$ yaw angle for range of angles of attack.", "timestamp": "2026-07-22T04:10:35.433687+00:00"}