PDF/ocr_dataset_1075.jsonl

{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 2, "total_pages": 30, "image_filename": "19930085975_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:44:35.306698+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 79, "total_pages": 114, "image_filename": "19930086061_p79.jpg", "text": "NACA RM L9J07\n75\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\nP -1\n0\n1\nRight semispan\n\nUpper\nLower\n\n120°\n\n(c) $\\Psi = 20^\\circ$\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\nP -1\n0\n1\nRight semispan\n\nUpper\nLower\n\n35°\n\n(d) $\\Psi = 35^\\circ$\n\nNACA\n\nFigure 26.- Concluded.", "timestamp": "2026-07-22T06:44:36.997967+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 12, "total_pages": 22, "image_filename": "19930085928_p12.jpg", "text": "```markdown\nNACA RM No. A9A31 CONFIDENTIAL 11\n\nWhen solved for Mach number, equation (A5) becomes\n\n$$\nM = \\left\\{ \\frac{2}{\\gamma-1} \\left[ \\left( \\frac{\\gamma+1}{2} \\right) e^{\\frac{1-\\gamma}{\\gamma} K \\left( \\frac{x}{L} \\right)} - 1 \\right] \\right\\}^{1/2} \\quad \\text{(A6)}\n$$\n\nThe relation between area and Mach number in an isentropic flow when $M_2 \\equiv M_{x/L=0} = 1.0$ is\n\n$$\n\\frac{A_2}{A} = M \\left\\{ \\frac{(\\gamma+1)/2}{1 + [(\\gamma-1)/2]M^2} \\right\\}^{2\\left(\\frac{\\gamma+1}{\\gamma-1}\\right)} \\quad \\text{(A7)}\n$$\n\nSubstituting equation (A6) in equation (A7),\n\n$$\n\\frac{A_2}{A} = \\left\\{ \\frac{2}{\\gamma-1} \\left[ \\left( \\frac{\\gamma+1}{2} \\right) e^{\\frac{1-\\gamma}{\\gamma} K \\left( \\frac{x}{L} \\right)} - 1 \\right] \\right\\}^{1/2} e^{\\frac{\\gamma+1}{2\\gamma} K \\left( \\frac{x}{L} \\right)} \\quad \\text{(A8)}\n$$\n\nREFERENCES\n\n1. Davis, Wallace F., and Edwards, Sherman S.: Experimental Investigation at Supersonic Speeds of Twin-Scoop Duct Inlets of Equal Area. III - Inlet Enclosing 37.2 Percent of the Maximum Circumference of the Forebody. NACA RM No. A8E04, 1948.\n\n2. Davis, Wallace F., Brajnikoff, George B., Goldstein, David L., and Spiegel, Joseph M.: An Experimental Investigation at Supersonic Speeds of Annular Duct Inlets Situated in a Region of Appreciable Boundary Layer. NACA RM No. A7G15, 1947.\n\n3. Evvard, John C., and Blakey, John W.: The Use of Perforated Inlets for Efficient Supersonic Diffusion. NACA RM No. E7C26, 1947.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:44:37.443241+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 83, "total_pages": 98, "image_filename": "19930086073_p83.jpg", "text": "NACA RM A9H04\n81\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\n| Angle of sideslip, $\\beta$, deg | | | | |\n| :--- | :--- | :--- | :--- | :--- |\n| 0.0 | 6.0 | 12.0 | 15.9 | |\n\n(b) $C_L$ vs $C_D$.\n\nFigure 18.— Continued.", "timestamp": "2026-07-22T06:44:37.895222+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 32, "total_pages": 33, "image_filename": "19930085900_p32.jpg", "text": "```markdown\nNACA RM L9D20\n31\n\nCONFIDENTIAL\n\n<!-- Image (229, 143, 769, 919) -->\n\nFigure 15.- Comparison of jets and strips simulating multiple steps.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:44:38.839411+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 33, "total_pages": 92, "image_filename": "19930085870_p33.jpg", "text": "34\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\nElliptical L.E. {CL, Cm}\nWedge L.E. {CL, Cm}\n\nCL\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\nCm\n.01\n0\n-.01\n\n[Figure: Graph showing CL and Cm vs alpha for Elliptical and Wedge L.E.]\n\nElliptical L.E. {CD, L/D}\nWedge L.E. {CD, L/D}\n\nCD\n.06\n.04\n.02\n0\n\nL/D\n6\n4\n2\n0\n\nalpha, deg\n-8 -6 -4 -2 0 2 4 6 8\n\nNACA\n\n(d) Wing 4. w=0.594; R=1,200,000.\nFigure 5. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:43.395183+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 46, "total_pages": 46, "image_filename": "19930085548_p46.jpg", "text": "NACA RM No. E8L30\n45\n\n10777\n\n[Figure: Graph showing Power vs. Manifold pressure with three curves: \"Power available, good scavenging efficiency\", \"Power required by compressor\", and \"Power available, poor scavenging efficiency\". NACA logo in bottom right corner of graph.]\n\nManifold pressure\n\nFigure 23. - Compressor and engine power relations in gas-generator engine.", "timestamp": "2026-07-22T06:44:44.758639+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 73, "total_pages": 78, "image_filename": "19930082483_p73.jpg", "text": "NACA TN No. 1807\n71\n\nNozzle and rotor-blading losses corrected to sea level, hp\n\nAdmission\n(deg)\n360\n270\n240\n180\n120\n90\n\nCorrected rotor speed, $N/\\sqrt{\\theta_1}$, rpm\n\n(e) Nozzle and rotor-blading losses.\n\nFigure 9. - Concluded. Calculated and experimental\nlosses for gas turbine operating at total-\npressure ratio of 2.0 for full and partial\nadmissions.", "timestamp": "2026-07-22T06:44:44.914879+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 15, "total_pages": 34, "image_filename": "19930085913_p15.jpg", "text": "```markdown\n14\n\nTABLE I.- EXPERIMENTAL DATA - Continued\n\n| Model | Run | $q_p$ (lb/sq ft) | $V_p$ (fps) | Mach number | Distance of weight from root (percent 1) | Frequencies (cps) | | | | Phase-angle relationship of bending and torsional stresses. (Ref indicates reference strain-gage trace) | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | **Natural** | | | **Flutter** | **2nd natural mode** | | | **3rd natural mode** | | | **Flutter mode** | | | |\n| | | | | | | 1st | 2nd | 3rd | | 1 (deg) | 2 (deg) | 3 (deg) | 4 (deg) | 1 (deg) | 2 (deg) | 3 (deg) | 4 (deg) | 1 (deg) | 2 (deg) | 3 (deg) | 4 (deg) |\n| Model A<br>Unswept untapered wing<br>Weight moved along midchord<br>line; $e_w = 0$<br>Reynolds number $\\frac{Vc}{\\nu}$ 3775.7$p$ | 14 | 17.23 | 129.1 | 0.1120 | 0 | 2.22 | 13.47 | 19.35 | 14.57 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 18 | 0 | 198 |\n| | 15 | 19.00 | 131.7 | .1135 | 5.55 | 2.30 | 13.33 | 19.48 | 14.63 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 26 | 0 | 226 |\n| | 16 | 28.38 | 161.2 | .1390 | 22.20 | 2.27 | 10.00 | 19.40 | $12.38$<br>$24.55$ | - | Ref | - | 180 | Ref | - | 0 | - | --- | --- | - | --- |\n| | 17 | 28.40 | 161.3 | .1390 | 27.77 | 2.24 | 8.95 | 18.92 | 10.81 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 4 |\n| | 18 | 29.65 | 164.9 | .1420 | 33.33 | 2.19 | 8.29 | 18.37 | 10.85 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 8 |\n| | 19 | 29.02 | 163.1 | .1405 | 38.90 | 2.06 | 8.03 | 17.80 | 10.60 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 25 |\n| | 20 | 27.40 | 158.5 | .1365 | 44.40 | 2.02 | 8.05 | 17.22 | 11.85 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 0 |\n| | 21 | 31.84 | 171.2 | .1475 | 50.00 | 1.89 | 8.43 | 16.69 | 8.00 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 180 |\n| | 22 | 32.33 | 172.5 | .1485 | 55.50 | 1.78 | 9.09 | 16.29 | 8.20 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 346 | - | 130 |\n| | 23 | 31.53 | 170.3 | .1466 | 61.11 | 1.67 | 10.07 | 15.99 | 8.63 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 0 | 0 | 0 |\n| | 24 | 32.45 | 169.0 | .1490 | 66.66 | 1.59 | 11.38 | 15.60 | 7.74 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 34 | - | 217 |\n| | 25 | 33.77 | 172.5 | .1520 | 72.20 | 1.49 | 12.76 | 15.36 | $7.23$<br>$28.90$ | - | Ref | - | 180 | Ref | - | 0 | - | --- | --- | - | --- |\n| | 26 | 33.76 | 172.7 | .1520 | 77.80 | 1.39 | 13.04 | 15.00 | $6.25$<br>$12.50$ | - | Ref | - | 180 | Ref | - | 0 | - | --- | --- | - | --- |\n| | 27 | 9.00 | 88.9 | .0780 | 83.30 | 1.30 | 12.96 | 14.81 | 13.14 | - | Ref | - | 180 | Ref | - | 0 | - | Ref | 73 | 0 | 240 |\n| | 28 | 10.06 | 94.1 | .0825 | 94.40 | ---- | ---- | ---- | 12.70 | - | --- | - | --- | --- | - | - | - | Ref | 34 | 0 | 221 |\n\n[Figure: Diagram of wing model showing dimensions l, 2b, t, and coordinate axes]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\nNACA RM 19F24\n```", "timestamp": "2026-07-22T06:44:47.329531+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 18, "total_pages": 22, "image_filename": "19930085922_p18.jpg", "text": "3C\nNACA RM No. L9C23\n\n$$ \\frac{pb}{2V} $$\n\n$$ \\delta_{a_2} $$\n(deg)\n$$ \\triangle - 4.4 $$\n$$ \\nabla - 1.8 $$\n$$ \\nabla \\quad 1.8 $$\n$$ \\square \\quad 4.4 $$\n\n[Figure: Graph plotting pb/2V against Mach number, M, with data points for various δa2 values. The NACA logo is visible in the bottom right corner of the graph.]\n\nMach number, M\n\nFigure 8.- Variation with Mach number of the wing-tip helix angle obtained for various aileron deflections; vertical fins on; $$ \\delta_{a_r} = 0 $$.\n\n17", "timestamp": "2026-07-22T06:44:47.710581+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 22, "total_pages": 52, "image_filename": "19930085911_p22.jpg", "text": "NACA RM E9F22\nCONFIDENTIAL\n21\n\n[Figure: A twin-engine propeller airplane in flight, viewed from below and slightly behind. The aircraft has U.S. military star insignia on the wings and tail. Markings include \"FQ-887\" on the fuselage and \"483887\" on the tail fin. A ram-jet unit is mounted beneath the right wing.]\n\nNACA\nC-22249\n9-20-48\n\nFigure 1. - Supersonic 16-inch ram-jet unit mounted beneath airplane wing.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:48.889377+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 47, "total_pages": 60, "image_filename": "19930085862_p47.jpg", "text": "```markdown\nNACA RM No. L9A07\n45\n\nFlap configuration\noff\n\nSplit\n\nDrooped-nose and\nsplit and fences\n\nLeading-edge\nand split\n\nLeading-edge and\nsplit and fences\n\nLeading-edge\nand fences\n\n<!-- Image (209, 135, 838, 882) -->\n\nFigure 15.- Effects of high-lift and stall-control devices on the aileron\nhinge-moment parameters $C_{h\\delta}$, $C_{h\\alpha}$, $P_{R\\delta}$, and $P_{R\\alpha}$.\n```", "timestamp": "2026-07-22T06:44:51.626311+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 5, "total_pages": 55, "image_filename": "19930085966_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM L9B17\n\n$\\theta_{t5}$\nratio of absolute total temperature at exhaust-nozzle exit\nto absolute static temperature at NACA standard atmospheric\nconditions at sea level, 519° F ($T_{t5}/519$)\n\n$\\tau_t$\nratio of absolute total temperature at exhaust-nozzle exit\nto absolute total temperature at combustion-chamber\ninlet ($T_{t5}/T_{t3}$)\n\n$F/\\delta_0$\nthrust reduced to NACA standard atmospheric conditions at\nsea level, pounds\n\n$\\frac{W_a}{\\delta_0}\\sqrt{\\theta_{t5}}$\nreduced air-flow parameter, pounds per second\n\n$\\frac{3600W_f n_b}{\\delta_0 \\frac{\\tau_t - 1}{\\tau_t} \\sqrt{\\theta_{t5}}}$\nreduced fuel-consumption parameter, pounds per hour\n\nSubscripts:\n0 to 7\nconditions at the corresponding stations indicated in figure 7\n\nx\npoint in any cross section\n\ny\npoint between stations 3 and 4\n\nDESCRIPTION OF APPARATUS\n\nFigure 1 shows the combustion-chamber shell to be a semicircular\nsection suspended from a flat, horizontal structure containing a built-\nup truss. The $\\frac{7}{32}$-inch space containing the truss served as a cooling\nshroud for the top of the combustion chamber. The cooling of the curved\nportion of the combustion-chamber walls was provided for by the addition\nof a shroud giving a cooling-air passage measuring 1 inch between inner\nand outer walls. The combustion-chamber length was increased to 5 feet\nto obtain more complete combustion at the higher fuel flows. The fuel\nlines were altered so that the individual lines from the burners were\nmanifolded outside of the combustion chamber, instead of inside, in\norder to simplify maintenance of the setup.\n\nThe burners used were modified versions of that shown in figure 2.\nFuel was brought into the burners by two lines, the pilot and main\nfeed lines, as shown in figure 2. The pilot fuel traveled through the\npilot distributor rake and was projected in 0.024-inch-diameter\nstreams against the interior walls of the upstream region of the pilot\nhousing. The fuel in liquid condition was ignited and burned from\nthe walls of the pilot housing. The pilot housing created a low-velocity\nregion, necessary for the ignition and existence of the pilot flame. The\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:51.983616+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 80, "total_pages": 114, "image_filename": "19930086061_p80.jpg", "text": "```markdown\n76\nNACA RM L9J07\n\n<!-- Image (57, 78, 922, 905) -->\n\nFigure 27.- Flow characteristics over wing 1 as indicated by surface tufts.\n```", "timestamp": "2026-07-22T06:45:02.619569+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 26, "total_pages": 29, "image_filename": "19930085899_p26.jpg", "text": "```markdown\nNACA RM No. 19A21\n\nDownwash angle, $\\epsilon$, deg\n6\n4\n2\n0\n-2\nM = 0.60\nM = 0.70\nM = 0.80\nM = 0.85\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\nSymbol $\\circ$ $\\triangle$ $\\square$ $\\nabla$ $\\diamond$ $\\triangleleft$ $\\circ$ $\\triangle$ $\\circ$ $\\diamond$\n\nDownwash angle, $\\epsilon$, deg\n6\n4\n2\n0\n-2\nM = 0.88\nM = 0.90\nM = 0.93\nM = 0.95\n[Figure: NACA logo]\n\n-80 -40 0 40 80\n-80 -40 0 40 80\n-80 -40 0 40 80\n-80 -40 0 40 80\nTail height, $h_t$, percent semispan\n\nFigure 11.- Effective downwash angles in region of tail plane for a model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil. Wing alone.\n\n25\n```", "timestamp": "2026-07-22T06:45:04.853489+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 84, "total_pages": 98, "image_filename": "19930086073_p84.jpg", "text": "82\nNACA RM A9H04\n\n<!-- Image (107, 126, 879, 804) -->\n\n(c) $C_L$ vs $C_m$.\n\nFigure 18.— Continued.", "timestamp": "2026-07-22T06:45:07.184433+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 13, "total_pages": 22, "image_filename": "19930085928_p13.jpg", "text": "12 CONFIDENTIAL NACA RM No. A9A31\n\n4. Kantrowitz, Arthur, and Donaldson, Coleman DuP.: Preliminary Investigation of Supersonic Diffusors. NACA ACR L5D20, 1945.\n\n5. Naumann: Wirkungsgrad von Diffusoren Bei Hohen Unterschallgeschwindigkeiten. FB Nr. 1705, Deutsche Luftfahrtforschung (Berlin-Aldershof), 1942.\n\n6. Ferri, Antonio, and Nucci, Louis M.: Preliminary Investigation of a New Type of Supersonic Inlet. NACA RM No. L6J31, 1946.\n\n7. Wyatt, DeMarquis D., and Hunczak, Henry R.: An Investigation of Convergent-Divergent Diffusors at Mach Number 1.85. NACA RM No. E6K21, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:07.219511+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 46, "total_pages": 149, "image_filename": "19930083192_p46.jpg", "text": "42\nNACA TN 1976\n\nPiloting and continuous rough air.- Previous analytical and experimental studies were made for a single gust isolated in space and with elevator fixed, whereas airplane flights are in continuous rough air with a pilot modifying the reactions of the airplane. Although there are few available data covering this condition, on the basis of statistical analysis, significant differences have been found to exist between pilots and between airplanes of the same type. The effect of different pilots was as much as 20 percent and differences between airplanes were 5 to 10 percent. It appears from these results that the pilot is more important than the variation between airplanes of the same type in determining the loads applied to the airplane. The variation due to both factors is estimated to average 15 percent.\n\nLittle or no correlation between the maximum angular motions of an airplane and the maximum loads in the same traverse has been found. It was found in another flight test that one pilot moved the controls five times as often as another pilot and no significant difference in the loads imposed was found. These results led to the conclusion that the effect of the pilot on the imposed load is due to his past actions and is not directly related to the effect of any single gust.\n\nAt the present time the question of the effect of piloting and airplane motion in continuous rough air on the prediction of gust load factors is only being approached. The redeeming feature at present is the fact that the evaluation of flight load experience under actual operating conditions will include piloting effects for the airplanes tested. The question still to be answered is whether data on piloting effects taken in the past on older types of airplanes are valid for modern high-speed airplanes flown by pilots with different training and with airplanes designed for different requirements of flying qualities and stability.\n\nUnsymmetrical gusts and airplane response.- Little is known concerning the response of airplanes under the action of a gust which strikes only one wing. The only material available is empirical in character and has been previously presented in this paper in connection with the studies of gust structure. In that material it was indicated that calculation by a simplified method, neglecting the alleviation effects of motion, would have to be adequate until more information concerning the significant parameters and the motions of the airplane is obtained.\n\nHorizontal tail loads.- The results given in figure 37 indicate that the static margin and the tail volume are not of primary importance in the determination of tail load. The data indicate a reasonably ordered variation of tail load with center-of-gravity position. The magnitude of the tail load varies rapidly with gradient distance, the variation ranging from about 50 percent for a gradient distance of 0 chords to 0 percent for a gradient distance of 16 chords. Inspection of", "timestamp": "2026-07-22T06:45:08.722153+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 33, "total_pages": 33, "image_filename": "19930085900_p33.jpg", "text": "32\nNACA RM L9D20\n\nCONFIDENTIAL\n\n<!-- Image (189, 156, 724, 920) -->\n\nFigure 16.- Comparison of strips simulating multiple steps.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:09.399717+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 13, "total_pages": 47, "image_filename": "19930085919_p13.jpg", "text": "12 CONFIDENTIAL NACA RM No. A9C21\n\nat which instability first occurred. The highest lift coefficients attained before the occurrence of longitudinal instability for the model with the various flaps are summarized in figure 19. The largest gain in this lift coefficient was produced by the split flap deflected $45^\\circ$ at the trailing edge of the wing. The addition of the leading-edge flaps of 50-percent wing span deflected $40^\\circ$ increased this lift coefficient slightly. However, with either of the leading-edge flaps of full wing span deflected $40^\\circ$, the elevon deflected $-20^\\circ$, and the split flap deflected $45^\\circ$, a lift coefficient greater than 1.0 was attained before longitudinal instability occurred.\n\nCONCLUSIONS\n\nFrom an experimental investigation at low speed of the effects of split flaps, elevons, and leading-edge devices on the characteristics of a wing-fuselage combination employing a wing swept back $63^\\circ$ it is concluded that:\n\n1. For the plain wing-fuselage combination, an increase of Reynolds number from 2.5 to 4.2 million increased the lift coefficient attained before the occurrence of longitudinal instability from about 0.4 to 0.5, but a further increase of Reynolds number to 7.2 million resulted in no improvement of this lift coefficient.\n\n2. The optimum chordwise position of the split flap for delaying the occurrence of longitudinal instability to a higher lift coefficient was the position with the flap hinge line coincident with the wing trailing edge. This was the only position of the split flap which greatly reduced the drag of the model at the higher lift coefficients.\n\n3. The rate of change of lift, pitching moment, and rolling moment with elevon deflection remained nearly constant up to an angle of attack of about $9^\\circ$, but decreased at greater angles of attack.\n\n4. The 50-percent-span leading-edge flaps gave no significant improvement in the pitching-moment characteristics of the model. However, the full-span leading-edge flaps deflected $40^\\circ$ increased the lift coefficient attained before the occurrence of longitudinal instability to a value greater than 1.0 with the split flap deflected $45^\\circ$ at the trailing edge of the wing and the elevon deflected $-20^\\circ$.\n\n5. The extended-nose flap of full wing span was about twice as effective as the drooped-nose flap of full wing span in reducing the\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:10.042514+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 3, "total_pages": 30, "image_filename": "19930085975_p3.jpg", "text": "NACA RM L9E10\nCONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEFFECTS OF MACH NUMBER AND SWEEP ON THE DAMPING-IN-ROLL\nCHARACTERISTICS OF WINGS OF ASPECT RATIO 4\n\nBy Richard E. Kuhn and Boyd C. Myers, II\n\nSUMMARY\n\nThe damping-in-roll characteristics of three wings with an aspect ratio of 4, a taper ratio of 0.6, sweep angles of $3.6^\\circ$, $32.6^\\circ$, and $46.7^\\circ$ at the quarter-chord line, and with the NACA 65A006 section have been determined through the Mach number range from 0.4 to 0.91 and angle-of-attack range from $0^\\circ$ to $6.5^\\circ$ in the Langley high-speed 7- by 10-foot tunnel by the free-roll method. The results indicated that the increase in magnitude of the damping-in-roll coefficient $C_{l_p}$ with Mach number and the decrease with sweep angle, at low angles of attack, agreed well with the theoretical variations. The damping coefficient increased markedly with angle of attack (in the test range) particularly at the higher Mach numbers investigated.\n\nINTRODUCTION\n\nLow-speed experimental data and theory (references 1 and 2) indicate an appreciable reduction in the damping-in-roll properties of a wing as the sweep angle is increased. The theoretical manner in which these effects are affected by compressibility is treated in references 2 and 3. Little experimental data, however, are available at high-subsonic Mach numbers for comparison with theory. Accordingly, an extensive investigation is being conducted in the Langley high-speed 7- by 10-foot tunnel to determine the effects of sweep angle and Mach number on the damping-in-roll characteristics of a series of wings. The first wing investigated was a $35^\\circ$ sweptback wing of aspect ratio 3, and the damping-in-roll characteristics of this wing at high-subsonic Mach numbers are presented in reference 4.\n\nThe present paper presents the results of an experimental determination of the damping-in-roll characteristics of three wings of aspect ratio 4 and taper ratio 0.6 with sweep angles of $3.6^\\circ$, $32.6^\\circ$, and $46.7^\\circ$ referred to the quarter-chord line. The investigation utilized the free-roll technique described in reference 4, and the tests were made at angles of attack of $0.30^\\circ$, $3.45^\\circ$, and $6.5^\\circ$ through a Mach number\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:15.340850+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 34, "total_pages": 92, "image_filename": "19930085870_p34.jpg", "text": "NACA RM No. L9D07\n35\n\nCONFIDENTIAL\n\n.24\n.16\n.08\n$C_L$\n0\n-.08\n-.16\n-.24\n\nElliptical L.E. $\\begin{cases} \\circ C_L \\\\ \\square C_m \\end{cases}$\nWedge L.E. $\\begin{cases} \\triangle C_L \\\\ \\diamond C_m \\end{cases}$\n\n.01\n$C_m$\n0\n-.01\n\n.06\n.04\n.02\n$C_D$\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8\n\nElliptical L.E. $\\begin{cases} \\circ C_D \\\\ \\square L/D \\end{cases}$\nWedge L.E. $\\begin{cases} \\triangle C_D \\\\ \\diamond L/D \\end{cases}$\n\n6\n$L/D$\n4\n2\n0\n\n$\\alpha$, deg\n\n(e) Wing 5. w=0.675; R=1,080,000.\nFigure 5. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:16.124826+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 3, "total_pages": 30, "image_filename": "19930085970_p3.jpg", "text": "NACA RM A9E09\nCONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING\nA WING SWEPT BACK 63°.- CHARACTERISTICS FOR\nSYMMETRICAL WING SECTIONS AT HIGH SUBSONIC\nAND MODERATE SUPERSONIC MACH NUMBERS\n\nBy Newton A. Mas\n\nSUMMARY\n\nResults of wind-tunnel tests are presented for a wing with the\nleading edge swept back 63° and of symmetrical section in combination\nwith a body at Mach numbers from 0.5 to 0.95 and from 1.09 to 1.51.\nThe test Reynolds numbers varied from 0.35 to 0.52 million. Measured\nlift, drag, and pitching-moment coefficients for the configuration\nare compared with corresponding calculated characteristics. The\nresults indicate that available analytical methods may be used with\nconfidence in the prediction of the variations with Mach number of\nthe lift of highly swept wings. It is also found that the measured\ntrends of the minimum drag coefficient with Mach number compare\nfavorably with those indicated by theory throughout the Mach number\nrange of the tests. The low Reynolds numbers of the tests virtually\ninvalidate any quantitative comparison of the measured characteristics\nof pitching moment and drag due to lift with those calculated by the\nmethods of inviscid theory. However, the results are useful in\nindicating gross changes with Mach number of the aerodynamic-center\nlocation and the approximate magnitude of the maximum lift-drag\nratio to be expected for a highly swept wing configuration at\nmoderate supersonic Mach numbers.\n\nINTRODUCTION\n\nR. T. Jones has indicated in reference 1 the possibility of\ndeveloping practicable values of maximum lift-drag ratio at super-\nsonic Mach numbers with wings swept well behind the Mach cones\nemanating from the leading edges. To examine this possibility experi-\nmentally and to determine the aerodynamic properties of such highly\nswept wings under other flight conditions, an extensive wind-tunnel\ninvestigation has been undertaken in several facilities of the Ames\nAeronautical Laboratory. Tests have been completed (references 2, 3,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:45:16.812423+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 11, "total_pages": 92, "image_filename": "19930085951_p11.jpg", "text": "NACA RM L9D29\nUNCLASSIFIED\n9\n\nreduction in blade-section thickness of only the outboard blade sections (from 8 to 5 percent at the 0.7 radius) resulted in gains in propeller efficiency up to 4 percent. These improvements in efficiency appear to be limited to the range of advance ratio for which the propeller is designed. It should be pointed out, however, that the helical-tip Mach number of the NACA propellers did not exceed 0.8 for the conditions of operation shown in figure 28.\n\nFigures 29 and 30 have been prepared to emphasize the importance of blade-section thickness in the design of propellers to operate at airspeeds where the tip Mach numbers are below the critical value. Figure 29 shows the effect of airspeed on the difference in efficiency between the NACA 10-(3)(08)-03 and NACA 10-(3)(12)-03 two-blade propellers when operating at a constant power coefficient of 0.15 and a constant rotational speed of 1140 rpm. The thinner blade sections of the NACA 10-(3)(08)-03 propeller effect an increase in efficiency of 10 percent with an increase in airspeed from 260 to 420 miles per hour. The corresponding change in helical-tip Mach number is from 0.63 to 0.76 as shown in figure 29. Figure 30 shows the effect of airspeed on the difference in efficiency between the NACA 10-(3)(05)-045 and NACA 10-(3)(08)-045 two-blade propellers when operating at a constant power coefficient of 0.15 and a constant rotational speed of 1140 rpm. The thinner outboard blade sections of the NACA 10-(3)(05)-045 propeller effect an increase in efficiency of 4 percent with an increase in airspeed up to 220 miles per hour, but from 220 to 460 miles per hour the beneficial effects of the thinner blade sections are gradually lost. Apparently the thicker outboard blade sections of the NACA 10-(3)(08)-045 propeller can carry their loads just as efficiently as the thinner sections of the NACA 10-(3)(05)-045 propeller at an airspeed of 460 miles per hour. The helical-tip Mach number at this airspeed is only 0.8, and it is possible that beneficial effects of the thinner blade sections may appear for different radial distributions of section Mach number. It should be realized that none of the NACA propellers was designed to operate at advance ratios as high as 3.2, otherwise higher efficiencies might be expected at the higher values of advance ratio because of a different pitch distribution. Since the superiority of the thinner outboard blade sections appears principally in the range of advance ratio for which the propeller was designed, substantial gains in efficiency through the use of thinner outboard blade sections may not be realized unless care is used in selecting the radial pitch distribution.\n\nEffect of thickness ratio and compressibility on propeller characteristics.— The effect of compressibility on the maximum efficiency of NACA propellers having different blade-section thicknesses is shown in figure 31 for a blade angle at the 0.75 radius of 45°. Figure 31(a) shows that the NACA 10-(3)(12)-03 propeller, which has the thickest outboard blade sections in the 0.03-solidity group, suffers the greatest efficiency losses at the higher tip Mach numbers. These losses begin for this propeller at a helical-tip Mach number of about 0.825, and the loss amounts to 26 percent at a helical-tip Mach number of 1.1. The loss in efficiency\n\nCONFIDENTIAL\nUNCLASSIFIED", "timestamp": "2026-07-22T06:45:17.773298+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 74, "total_pages": 78, "image_filename": "19930082483_p74.jpg", "text": "72\nNACA TN No. 1807\n\nAdmission, deg\n360 180 120\n\nPower corrected to sea level, hp/lb driving fluid\n30\n25\n20\n15\n10\n5\n0\n\nNet power\nRotor-tip leakage loss\nBearing loss\nPumping loss\nDriving-fluid losses\nNozzle and rotor-blade losses\n\n2 6 10 14x10³ 2 6 10 14x10³ 2 6 10 14x10³\nCorrected rotor speed, $N/\\sqrt{\\theta_j}$, rpm\n\nNACA\n\nFigure 10. - Specific ideal power, specific net power, and applicable specific power losses for 360° (full), 180°, and 120° admissions. Total-pressure ratio, 2.0.", "timestamp": "2026-07-22T06:45:19.090010+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 19, "total_pages": 22, "image_filename": "19930085922_p19.jpg", "text": "18\nNACA RM No. L9C23\n\n<!-- Image (152, 138, 751, 469) -->\n\n<!-- Image (136, 526, 762, 873) -->\n\nFigure 9.- Variation of the aileron effectiveness with Mach number for various angles of attack; vertical fins off.", "timestamp": "2026-07-22T06:45:19.634750+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 1, "total_pages": 25, "image_filename": "19930085979_p1.jpg", "text": "```markdown\nFILE COPY\nNo 8\n\nRESTRICTED\n\nCopy\nRM E9E12\n\n428\n\nNACA RM E9E12\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF HOT-GAS BLEEDBACK FOR ICE\nPROTECTION OF TURBOJET ENGINES\n\nIII - NACELLE WITH SHORT STRAIGHT AIR INLET\n\nBy Robert S. Ruggeri and Edmund E. Callaghan\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\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 AS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H ST., N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, 50 U.S.C. 31 and 32. The 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, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\n\nAUTHORITY J. W. CROWLEY\nDATE 12-14-53 CHANGE #1909 E.L.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nAugust 4, 1949\n\nRESTRICTED\n```", "timestamp": "2026-07-22T06:45:20.067474+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 23, "total_pages": 52, "image_filename": "19930085911_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:45:22.778138+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 48, "total_pages": 60, "image_filename": "19930085862_p48.jpg", "text": "46\nNACA RM No. L9A07\n\n<!-- Image (119, 110, 832, 874) -->\n\nFigure 16.- Effect of various amounts of aerodynamic balance on aileron hinge-moment parameter $C'_{h\\delta}$.", "timestamp": "2026-07-22T06:45:26.783086+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 55, "total_pages": 62, "image_filename": "19930082918_p55.jpg", "text": "66\nNACA TN 1940\n\n<!-- Image (209, 110, 818, 997) -->\n\nFigure 15.- Effect of aging on secondary creep rates at 60,000 psi and 1200° F of low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:45:30.662290+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 49, "total_pages": 72, "image_filename": "19930085491_p49.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:45:37.060535+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 2, "total_pages": 25, "image_filename": "19930085979_p2.jpg", "text": "NACA RM E5E12\nRESTRICTED\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION OF HOT-GAS BLEEDBACK FOR ICE\nPROTECTION OF TURBOJET ENGINES\n\nIII - NACELLE WITH SHORT STRAIGHT AIR INLET\n\nBy Robert S. Ruggeri and Edmund E. Callaghan\n\nSUMMARY\n\nAerodynamic and icing investigations have been conducted in\nthe NACA Lewis icing research tunnel on a two-thirds-scale model\nof a turbojet-engine nacelle with a short straight air inlet. An\ninvestigation of a hot-gas bleedback system consisting of several\norifices peripherally located around the inlet was conducted for\nboth dry-air and icing conditions.\n\nThe most uniform temperature distribution was obtained at a\nbleedback of 4.9 percent for a plenum-chamber-gas temperature of\n1000° F and yielded an average model-dry-air-temperature rise of\n50° F with a maximum deviation from the average model-air-\ntemperature rise of 10° F. For this condition, icing protection\nwas afforded for the inlet screen but not for the accessory housing.\nSatisfactory agreement between calculated and measured heat\nrequirements was obtained for the icing conditions investigated.\n\nINTRODUCTION\n\nAs part of a general program to provide icing protection for\nturbojet engines, a nacelle with several air inlets was experimen-\ntally investigated in the icing research tunnel of the NACA Lewis\nlaboratory to establish a reasonable design criterion for hot-gas\nbleedback systems. This investigation is a continuation of the\ngeneral program outlined in reference 1 and was conducted with the\nsame turbojet-engine nacelle, but with a short straight air inlet.\n\nThis short air inlet was used to determine whether sufficient\njet mixing would take place to make hot-gas bleedback feasible as a\nsystem of ice prevention for this type of installation. The nacelle\nwas two-thirds full scale and the model was provided with orifices\n\nRESTRICTED", "timestamp": "2026-07-22T06:45:57.564222+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 16, "total_pages": 34, "image_filename": "19930085913_p16.jpg", "text": "```markdown\nTABLE I.- EXPERIMENTAL DATA - Continued\n\nNACA RM L9F24\n\n| Model | Run | $q_F$ (lb/sq ft) | $V_F$ (fps) | Mach number | Distance of weight from root (percent l) | Frequencies (cps) | | | | Phase-angle relationship of bending and torsional stresses. (Ref indicates reference strain-gage trace) | | | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | **Natural** | | | **Flutter** | **2nd natural mode** | | | | **3rd natural mode** | | | | **Flutter mode** | | | |\n| | | | | | | 1st | 2nd | 3rd | | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) |\n| Model B<br>Swept untapered wing;<br>$\\Lambda = 45^\\circ$<br>Weight moved along leading<br>edge; $e_w = -1$<br>Reynolds number $\\frac{qc}{\\mu} 5310.2V_F$ | 29 | 29.42 | 162.0 | 0.1419 | -11.11 | 2.47 | 14.91 | 20.30 | 11.12 | Ref | - | - | 0 | Ref | 0 | 0 | 180 | Ref | 24 | 0 | 24 |\n| | 30 | 29.29 | 161.8 | .1416 | 0 | 2.48 | 14.91 | 20.23 | 11.28 | Ref | - | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 31 | 27.22 | 156.2 | .1365 | 5.55 | 2.48 | 14.94 | 18.87 | 11.38 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 32 | 25.30 | 150.6 | .1315 | 11.11 | 2.48 | 12.81 | 17.20 | 10.24 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 11 | 0 | 11 |\n| | 33 | 22.13 | 140.9 | .1229 | 16.66 | 2.46 | 10.51 | 17.24 | 8.96 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 34 | 20.51 | 135.7 | .1183 | 22.20 | 2.44 | 8.82 | 17.59 | 7.85 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 35 | 21.10 | 137.7 | .1200 | 27.77 | 2.41 | 7.75 | 17.89 | 6.79 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 36 | 27.84 | 158.4 | .1380 | 33.33 | 2.38 | 7.09 | 18.02 | 6.25 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 24 | 0 | 24 |\n| | 37 | 53.17 | 219.8 | .1918 | 38.90 | 2.27 | 6.76 | 18.00 | 5.96 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 162 | 0 | 162 |\n| | 38 | 99.90 | 296.5 | .2585 | 44.40 | 2.16 | 6.75 | 17.83 | $^a$16.35 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 39 | 114.50 | 327.9 | .2855 | 50.00 | 2.04 | 6.96 | 17.54 | $^a$16.96 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | --- |\n| | 40 | 143.50 | 364.8 | .3213 | 55.50 | 1.91 | 7.35 | 17.21 | $^a$18.47 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 12 | 12 | 0 |\n| | 41 | 155.80 | 382.7 | .3357 | 61.11 | 1.78 | 7.94 | 16.88 | $^a$28.6 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Poor record | | | |\n| | 42 | 159.60 | 389.1 | .3400 | 66.66 | 1.69 | 8.73 | 16.42 | $^a$27.60<br>$^a$41.52 | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 180 |\n| | 43 | 161.20 | 392.7 | .3421 | 72.20 | 1.54 | 9.52 | 15.84 | $^a$21.20<br>$^a$44.30 | Ref | 0 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 0 | -- | 180 |\n| | 44 | 75.18 | 264.2 | .2290 | 77.80 | 1.41 | 10.30 | 15.26 | 15.88 | Ref | 0 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 45 | 33.89 | 176.4 | .1525 | 83.30 | 1.33 | 10.74 | 14.78 | 13.70 | Ref | 0 | 0 | 0 | No record | | | | Ref | 0 | 0 | 0 |\n| | 46 | 14.16 | 113.5 | .0980 | 88.90 | 1.24 | 10.91 | 14.11 | 13.13<br>$^a$39.0 | Ref | 0 | 0 | --- | Ref | - | 0 | 0 | --- | --- | -- | --- |\n| | 47 | 9.55 | 93.3 | .0805 | 94.40 | 1.16 | 10.48 | 13.90 | $^a$12.45<br>$^a$77.40 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 | --- | --- | -- | --- |\n| | 48 | 8.52 | 88.1 | .0760 | 96.60 | 1.10 | 9.92 | 13.63 | 12.50 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 | Ref | 4 | 0 | 328 |\n\n[Figure: Diagram of wing model showing length l, thickness t, and width 2b]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\n15\n```", "timestamp": "2026-07-22T06:46:02.034558+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 12, "total_pages": 92, "image_filename": "19930085951_p12.jpg", "text": "```markdown\n10\nUNCLASSIFIED\nNACA RM L9D29\n\ndue to compressibility is more gradual for the NACA 10-(3)(08)-03R pro-\npeller, and the serious losses do not begin until a helical-tip Mach number\nof about 0.875 is reached. For this propeller the loss in efficiency due\nto compressibility amounts to about 19 percent at a helical-tip Mach number\nof 1.1. The maximum efficiency of the NACA 10-(3)(08)-03 propeller, which\nhas the thinnest blade sections in the 0.03-solidity group, is about\n2 percent higher than the maximum efficiency of the other two propellers\nin the range of helical-tip Mach numbers below the critical value. The\ncritical value of tip Mach number is perhaps slightly higher for the\nNACA 10-(3)(08)-03 propeller than for the NACA 10-(3)(08)-03R propeller,\nand the loss in maximum efficiency due to compressibility amounts to about\n16 percent at a helical-tip Mach number of 1.1.\n\nFigure 31(a) shows that a reduction in blade-section thickness from\n12 to 8 percent at the 0.7 radius, or approximately one-third all along\nthe radius, resulted in a gain in propeller efficiency of about 12 percent\nat a helical-tip Mach number of 1.1. At this same helical-tip Mach number\na reduction in blade-section thickness of only the inboard blade-sections\n(from 30 to 13 percent at the 0.3 radius) resulted in a gain in propeller\nefficiency of about 5 percent.\n\nFigure 31(b) shows the effect of compressibility on the maximum effi-\nciency of the NACA propellers in the 0.045-solidity group. Again, the\npropeller having the thickest outboard blade sections (NACA 10-(3)(08)-045)\nsuffers the greatest efficiency losses at the higher tip Mach numbers.\nFrom a helical-tip Mach number of 0.90 to 1.1 the efficiency loss due to\ncompressibility for this propeller amounts to 18 percent. Over this same\nrange of helical-tip Mach number the propeller having the thinnest outboard\nblade sections (NACA 10-(3)(05)-045) has a loss in efficiency due to com-\npressibility of only 9 percent. The NACA 10-(3)(062)-045 propeller, which\nhas the thickest shank sections, shows a loss in efficiency due to com-\npressibility of 13 percent at a helical-tip Mach number of 1.1. These\nlosses at the higher tip Mach numbers are more gradual for the\nNACA 10-(3)(062)-045 propeller than for the NACA 10-(3)(08)-045 propeller,\nwhich has thinner blade sections at the shank but thicker outboard blade\nsections. The critical values of helical-tip Mach numbers are approxi-\nmately the same for all the propellers in the 0.045-solidity group, and\nthe differences in maximum efficiency are small at tip Mach numbers below\nthe critical value. The maximum efficiency of the NACA 10-(3)(062)-045A\npropeller is the same as for the thinner NACA 10-(3)(05)-045 propeller\nexcept at the lower values of helical-tip Mach number, where the thinner\nblade perhaps has a slight advantage.\n\nThe curves in figure 31(b) show that a reduction in blade-section\nthickness of only the outboard blade sections (from 8 to 5 percent at the\n0.7 radius) resulted in a gain in propeller efficiency of about 12 percent\nat a helical-tip Mach number of 1.1. At this same helical-tip Mach number\na reduction in blade-section thickness of only the inboard blade sections\n(from 16.6 to 13 percent at the 0.3 radius) resulted in a gain in propeller\n\nUNCLASSIFIED\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:46:03.660209+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 56, "total_pages": 62, "image_filename": "19930082918_p56.jpg", "text": "NACA TN 1940\n67\n\n<!-- Image (68, 110, 902, 993) -->\n\nFigure 16.- Effect of aging at 1400° F on 1200° F rupture strength of low-carbon M-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:46:06.513598+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 49, "total_pages": 60, "image_filename": "19930085862_p49.jpg", "text": "NACA RM No. L9A07\n47\n\n1.6\n1.2\n.8\n.4\n0 8 16 24\n$\\alpha$, deg\n$C_L$\n\n1.6\n1.2\n.8\n.4\n0 8 16 24\n$\\alpha$, deg\n$C_L$\n\n1.6\n1.2\n.8\n.4\n0 8 16 24\n$\\alpha$, deg\n$C_L$\n\n[Legend]\nCross flow\nRough\nIntermittently stalled\nCompletely stalled\n\n$C_L = 0.27$\n$\\alpha = 4.4^\\circ$\n\n$C_L = 0.95$\n$\\alpha = 9.0^\\circ$\n\n$C_L = 0.88$\n$\\alpha = 9.0^\\circ$\n\n$C_L = 0.56$\n$\\alpha = 8.6^\\circ$\n\n$C_L = 1.14$\n$\\alpha = 13.2^\\circ$\n\n$C_L = 1.04$\n$\\alpha = 13.1^\\circ$\n\n$C_L = 0.74$\n$\\alpha = 12.8^\\circ$\n\n$C_L = 1.25$\n$\\alpha = 17.3^\\circ$\n\n$C_L = 1.12$\n$\\alpha = 15.2^\\circ$\n\n$C_L = 0.79$\n$\\alpha = 14.9^\\circ$\n\n$C_L = 1.28$\n$\\alpha = 19.4^\\circ$\n\n$C_L = 1.26$\n$\\alpha = 19.3^\\circ$\n\n$C_L = 0.83$\n$\\alpha = 16.9^\\circ$\n\n$C_L = 1.29$\n$\\alpha = 21.4^\\circ$\n\n$C_L = 1.24$\n$\\alpha = 22.3^\\circ$\n\n[NACA logo]\n\n(a) Flaps off.\n(b) Extensible leading-edge flaps and split flaps.\n(c) Drooped-nose flaps and split flaps.\n\nFigure 17.— Stalling characteristics of $42^\\circ$ sweptback wing.", "timestamp": "2026-07-22T06:46:09.625482+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 14, "total_pages": 47, "image_filename": "19930085919_p14.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 13\n\ndrag of the model at the higher lift coefficients.\n\n6. With the full-span extended-nose flap deflected $40^\\circ$, the split flap deflected $45^\\circ$ at the trailing edge of the wing, and a wing loading of 40 pounds per square foot, a sinking speed of 30 feet per second was indicated for a lift coefficient of 1.0.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nAPPENDIX\n\nWIND-TUNNEL-WALL CORRECTIONS\n\nWind-tunnel-wall corrections for unswept reflection-plane models mounted on a 7-foot wall of a 7- by 10-foot wind tunnel have been presented in reference 6. For the purpose of the present report, the method used in reference 6 was modified to include the effects of sweepback upon the tunnel-wall corrections for a reflection-plane model mounted on a 10-foot wall.\n\nThe spanwise distribution of load was approximated by using two staggered horseshoe vortices as shown in figure 20. The normal method of summing the induced velocities of a doubly infinite image pattern was then followed. The induced velocity at the point P was computed separately for each horseshoe vortex and added in the following manner:\n\n$$\n\\left[ \\left( \\frac{w}{\\Gamma} \\right)_2 \\right]_{x=x'} + \\left[ \\left( \\frac{w}{\\Gamma} \\right)_2 \\right]_{x=x''} - \\left[ \\left( \\frac{w}{\\Gamma} \\right)_1 \\right]_{x=x''} = \\left( \\frac{w}{\\Gamma} \\right)_P\n$$\n\n(The subscript 1 refers to the horseshoe vortex having the trailing vortex a distance $y_1'$ from the plane of symmetry, and the subscript 2 refers to the horseshoe vortex having the trailing vortex a distance $y_1''$ from the plane of symmetry (fig. 20).)\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:46:12.279441+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930085983_p1.jpg", "text": "NACA RM A9I27\n\nCONFIDENTIAL\n\nCopy\nRM A9I27\n\nNACA\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING\nA WING SWEPT BACK $63^\\circ$.- EFFECTS AT SUBSONIC SPEEDS OF A\nCONSTANT-CHORD ELEVON ON A WING CAMBERED AND TWISTED\nFOR A UNIFORM LOAD AT A LIFT COEFFICIENT OF 0.25\n\nBy J. Lloyd Jones and Fred A. Demele\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\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\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nDecember 5, 1949\n\nCONFIDENTIAL\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 120\nEFFECTIVE DATE: SEPTEMBER 13, 1957\nWHL", "timestamp": "2026-07-22T06:46:12.476574+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 35, "total_pages": 92, "image_filename": "19930085870_p35.jpg", "text": "```markdown\n36\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n<!-- Image (66, 110, 916, 904) -->\n\n(f) Wing 6.w=0.731; R = 1,000,000.\nFigure 5. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:46:12.476820+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 27, "total_pages": 29, "image_filename": "19930085899_p27.jpg", "text": "```markdown\n26\n\nDownwash angle, $\\epsilon$, deg\n\nM = 0.98\n\nM = 1.00\n\nM = 1.03\n\nM = 1.05\n\n$\\alpha$, deg -2, -1, 0, 1, 2, 3, 4, 6, 8, 10\n\nSymbol $\\circ$ $\\triangle$ $\\square$ $\\nabla$ $\\diamond$ $\\triangleleft$ $\\triangleright$ $\\blacktriangleleft$ $\\blacktriangleright$ $\\blacklozenge$\n\nDownwash angle, $\\epsilon$, deg\n\nM = 1.08\n\nM = 1.10\n\nM = 1.15\n\nM = 1.18\n\n[Figure: NACA logo]\n\n-.80 -.40 0 .40 .80\n\n-.80 -.40 0 .40 .80\n\n-.80 -.40 0 .40 .80\n\n-.80 -.40 0 .40 .80\n\nTail height, $h_t$, percent semispan\n\nFigure 11.— Concluded.\n\nNACA RM No. 19A21\n```", "timestamp": "2026-07-22T06:46:15.538561+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 85, "total_pages": 98, "image_filename": "19930086073_p85.jpg", "text": "NACA RM A56H04\n\nLift coefficient, $C_L$\n\n$\\beta$, deg\n○ 0.0\n□ 6.0\n◇ 12.0\n△ 15.9\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 18.— Concluded.\n\nNACA\n\n83", "timestamp": "2026-07-22T06:46:19.191884+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 20, "total_pages": 22, "image_filename": "19930085922_p20.jpg", "text": "NACA RM No. L9C23\n19\n\n$$C_{l_\\delta}$$\n.003\n.002\n.001\n0\n.4 .5 .6 .7 .8 .9 1.0\nMach number, M\nFins on\nFins off\n\n$$\\left(\\frac{pb}{2V}\\right)_\\delta$$\n.008\n.006\n.004\n.002\n0\n.4 .5 .6 .7 .8 .9 1.0\nMach number, M\nFins on\nFins off\nNACA\n\nFigure 10.- Effect of the vertical fins on the aileron effectiveness;\n$\\alpha = 0.30^\\circ$.", "timestamp": "2026-07-22T06:46:19.645263+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 14, "total_pages": 22, "image_filename": "19930085928_p14.jpg", "text": "NACA RM No. A9A31\n\nCONFIDENTIAL\n\nSTATION 1\nDUCT ENTRANCE\n\nSTATION 2\nDUCT THROAT\n\nSTATION 3\nSETTLING CHAMBER\n\nSTATION 4\nVARIABLE EXIT\nTHROAT\n\nSLOT\n\n7.260 RAD.\n2.278\n1.352\n3.350\n3.000\n\n1.451\n.726\n.851\n1.500\n\nENTRANCE AREA, $A_1 = 0.2209$\nALL DIMENSIONS IN INCHES\n\nMODEL A\n\n.375 DIA.\n1.250 DIA.\n\nDIFFUSOR STATIONS (REFER TO FIGURE 2.)\nA B C D E F G H I J\n\nRAMP ANGLE\n\nL\n\nPITOT TUBE\n\nRAMP\n\nMODEL B\n\nNACA\n\n.400\n.726\n\nSLOT SIZES TESTED\n\nMODEL A\nHEIGHT LENGTH\n.085 .300\n.085 .450\n\nMODEL B\nHEIGHT LENGTH\n.085 .300\n.085 .450\n.044 .450\n\nFIGURE 1. - MODEL DIMENSIONS.\n\nCONFIDENTIAL\n\n13", "timestamp": "2026-07-22T06:46:22.416048+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 4, "total_pages": 30, "image_filename": "19930085970_p4.jpg", "text": "2 CONFIDENTIAL NACA RM A9E09\n\nand 4) and others are in progress to determine the principal aerodynamic characteristics of a configuration suggested by the analysis of reference 1 over a broad range of Mach and Reynolds numbers. The present investigation was initiated in the Ames 1- by 3-1/2-foot high-speed wind tunnel to determine the respective variations with Mach number of the lift, drag, and pitching-moment coefficients of the selected configuration at transonic Mach numbers beyond the reach of other currently available wind tunnels.\n\nThe configuration consists of a wing with the leading edge swept back $63^\\circ$ in combination with a body designed to have the minimum drag at supersonic speeds for a given length and volume. The wing was designed from aerodynamic considerations and from the structural criterion of reference 1 to provide useful maximum lift-drag ratios at Mach numbers up to about 1.5.\n\nThe results of the investigation are of additional value as an indication of the applicability of a number of linearized theories at Mach numbers for which they have been considered invalid.\n\nSYMBOLS\n\n| Symbol | Definition |\n| :--- | :--- |\n| A | aspect ratio $\\left(\\frac{b^2}{S}\\right)$ |\n| b | wing span, feet |\n| c | local chord, feet |\n| $\\bar{c}$ | mean aerodynamic chord $\\left(\\frac{\\int_0^{b/2} c^2 \\, dy}{\\int_0^{b/2} c \\, dy}\\right)$, feet |\n| $C_D$ | drag coefficient $\\left(\\frac{\\text{drag}}{qS}\\right)$ |\n| $C_{D_{\\text{min}}}$ | minimum drag coefficient |\n| $\\Delta C_D$ | increment in drag coefficient ($C_D - C_{D_{\\text{min}}}$) |\n| $C_L$ | lift coefficient $\\left(\\frac{\\text{lift}}{qS}\\right)$ |\n| $C_{L_{C_{D_{\\text{min}}}}}$ | lift coefficient corresponding to minimum drag coefficient |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:46:24.182860+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 50, "total_pages": 72, "image_filename": "19930085491_p50.jpg", "text": "NACA RM No. A8J04\n\nCONFIDENTIAL\n\nMach line\nFuselage outline\nLine of constant\npressure coefficient, P\nTheoretical line of\nlaminar boundary layer\nseparation\n\n[Figure: Diagram of a swept-back wing with contour lines labeled 0.0, -1.0, -2.0, -2.5, -3.0, and various line types as defined in the legend.]\n\nNACA\n\nFigure 6.- Location of theoretical line of laminar boundary-layer separation on 63° swept-back wing at zero angle of attack\n\nCONFIDENTIAL\n\n49", "timestamp": "2026-07-22T06:46:31.646498+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 75, "total_pages": 78, "image_filename": "19930082483_p75.jpg", "text": "1032\n\nNACA TN NO. 1807\n\nOver-all turbine efficiency, $\\eta$\n\nPercentage full-power output\n\nAdmission (deg)\n90\n120\n150\n180\n240\n300\n360\n\n[Figure: NACA logo]\n\nFigure 11. - Variation of over-all efficiency with power reduction by means of partial admission. Inlet total pressure, 45 inches mercury absolute; inlet total temperature, 800° R; corrected rotor speed, 8650 rpm; total-pressure ratio, 2.0.\n\n73", "timestamp": "2026-07-22T06:46:32.257970+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 21, "total_pages": 22, "image_filename": "19930085922_p21.jpg", "text": "20\n\nCoefficient of damping in roll, $C_{l_p}$\n\nUnstable\nStable\n\nTheory, ref. 3; $\\alpha = 0^\\circ$\n\n$\\alpha$\n$1.80^\\circ$\n$0.30^\\circ$\n$3.45^\\circ$\n$5.00^\\circ$\n$6.50^\\circ$\n\nNACA\n\nMach number, M\n\nFigure 11.- Variation of the coefficient of damping in roll $C_{l_p}$ with Mach number for various angles of attack; vertical fins off.\n\nNACA RM No. 19023", "timestamp": "2026-07-22T06:46:42.925475+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 47, "total_pages": 149, "image_filename": "19930083192_p47.jpg", "text": "NACA TN 1976\n43\n\ntables XII(a), XII(b), and XII(c) indicates that the decrease is due mainly to the increasing importance of the effect of pitch-angle increment as the gradient distance is increased and due to the fact that the total tail load is composed of one large term due to the gust, to which is added many small contributions from various sources which may be of either sign. The net result is that the tail load is, in many cases, small and equal to some of the small components. The calculation of the tail load thus requires an extremely high degree of accuracy in order to obtain an accurate answer.\n\nAs previously noted, the material presented in figure 37 and table XII shows the tail load at the time of maximum airplane acceleration. It is thought that the maximum tail load may result from a different sized gust than that which is critical for the wing or it may occur at a different time that at maximum acceleration. In view of the many uncertainties in the calculation of tail loads due to gusts, it has been suggested from time to time that the tail load be calculated as the increase in lift due to the gust alone multiplied by the factor $1 - \\frac{d\\epsilon}{d\\alpha}$. A slightly different concept would be to assume the gust velocity equal to about half that on the wing.\n\nWhat few experimental data from flight tests are available appear to be in essential agreement with the preceding suggestion. Table XV indicates that the ratio of the effective gust velocity on the tail as determined from tail-load measurements to the effective gust velocity for the wing is equal to 0.53 and 0.64 for the XB-15 and Q-2H airplanes, respectively. The results compare favorably with an estimated value for $1 - \\frac{d\\epsilon}{d\\alpha}$ of 0.5.\n\nIn the case of the canard airplane, inspection of the data in table XIV indicates that for a gust with a gradient distance of 8 chords the effects of pitch and vertical motion cancel so that the total stabilizer load is approximately equal to the load that would be imposed on the tail surface due to the gust alone. For the longer gradient distance of 17.5 chords, the cancellation is not complete, and the total tail load amounts to about two-thirds of the tail load resulting from the action of the gust. For the practical case, calculations of the horizontal tail load for canard airplanes might be based on the lift due to the change in angle of attack resulting from the gust if both the vertical motion and pitching motion of the airplane are neglected.\n\nIn summary, the available results indicate that it is not possible by means of detailed calculations to obtain an accurate estimate of the horizontal tail load due to the action of the gust. The tail load on", "timestamp": "2026-07-22T06:46:44.126019+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 57, "total_pages": 62, "image_filename": "19930082918_p57.jpg", "text": "68\nNACA TN 1940\n\n<!-- Image (125, 109, 882, 936) -->\n\nFigure 17.- Maximum true strain at rupture of low-carbon M-155 alloy aged at 1400° F, solution-treated 10 hours at 2200° F, and water-quenched.", "timestamp": "2026-07-22T06:46:46.841686+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 15, "total_pages": 22, "image_filename": "19930085928_p15.jpg", "text": "14\nCONFIDENTIAL\nNACA RM No. A9A31\n\n<!-- Image (100, 119, 867, 483) -->\n\nAll dimensions in inches\n\n| STATION | x | $A_2/A$ | STATION | x | $A_2/A$ |\n| :--- | :--- | :--- | :--- | :--- | :--- |\n| A | -0.400 | 0.914 | A | -0.400 | 0.748 |\n| B | 0 | 1.000 | B | 0 | 1.000 |\n| C | 0.375 | 0.994 | C | 0.375 | 0.994 |\n| D | 0.625 | 0.983 | D | 0.625 | 0.983 |\n| E | 1.125 | 0.936 | E | 1.125 | 0.936 |\n| F | 1.625 | 0.839 | F | 1.625 | 0.839 |\n| G | 2.125 | 0.648 | G | 2.125 | 0.648 |\n| H | 2.325 | 0.516 | H | 2.325 | 0.516 |\n| I | 2.500 | 0.334 | I | 2.500 | 0.305 |\n| J | 2.800 | 0.235 | J | 2.800 | 0.175 |\n\nModel A\nModel B\n[NACA logo]\n\nFigure 2.—Internal shape and areas of the model ducts.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:46:52.163977+00:00"}