PDF/ocr_dataset_1074.jsonl

{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 14, "total_pages": 22, "image_filename": "19930085922_p14.jpg", "text": "$5 \\times 10^6$\n\nReynolds number, $R$\n\nMach number, $M$\n\nNACA\n\nFigure 4.- Variation of test Reynolds number with Mach numbers based on the mean geometric chord of 1.046 feet.\n\nNACA RM No. L9C23\n\n13", "timestamp": "2026-07-22T06:42:24.362298+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 13, "total_pages": 34, "image_filename": "19930085913_p13.jpg", "text": "12\nNACA RM L9F24\n\nWings carrying a single weight (14 percent heavier than the wing) at a series of spanwise positions on the leading edge and on the midchord line were tested at sweepback angles of $0^\\circ$, $45^\\circ$, and $60^\\circ$. A comparison of the results obtained from the tests in which the wings were weighted at the leading edge indicated that the flutter speed was greatly affected by variation of spanwise weight position, and in these cases the change in sweepback angle had a small effect on the flutter speed. For the wings weighted at the midchord, the general effect of increase in sweepback was to increase the flutter speed and, as the sweep angle was increased, the effect of spanwise weight position became more pronounced.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Runyan, Harry L., and Watkins, Charles E: Flutter of a Uniform Wing with an Arbitrarily Placed Mass According to a Differential-Equation Analysis and a Comparison with Experiment. NACA TN 1848, 1949.\n\n2. Woolston, Donald S., and Runyan, Harry L.: Appraisal of Method of Flutter Analysis Based on Chosen Modes by Comparison with Experiment for Cases of Large Mass Coupling. NACA TN 1902, 1949.\n\n3. Runyan, Harry L., and Sewall, John L.: Experimental Investigation of the Effects of Concentrated Weights on Flutter Characteristics of a Straight Cantilever Wing. NACA TN 1594, 1948.\n\n4. Barmby, J. G., Cunningham, H. J., and Garrick, I. E.: Investigation of the Effects of Sweep on the Flutter of Cantilever Wings. NACA RM L8H30, 1948.", "timestamp": "2026-07-22T06:42:25.739062+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 79, "total_pages": 98, "image_filename": "19930086073_p79.jpg", "text": "```markdown\nNACA RM A53H04\n\nLift coefficient, $C_L$\n\nDrag coefficient, $C_D$\n\n| Symbol | Angle of sideslip, $\\beta$, deg |\n| :---: | :---: |\n| $\\circ$ | 0.0 |\n| $\\square$ | 6.0 |\n| $\\diamond$ | 12.0 |\n| $\\triangle$ | 15.9 |\n\n(b) $C_L$ vs $C_m$.\n\nFigure 17.— Continued.\n\n[Figure: NACA logo]\n\n77\n```", "timestamp": "2026-07-22T06:42:27.315265+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 10, "total_pages": 22, "image_filename": "19930085928_p10.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 9\n\nuseful in increasing the pressure recovery but they also improve the stability of operation of an air-induction system. In the operating range of some inlets, a large decrease in pressure recovery can result from a small transient increase in the mass flow to an engine. The thrust force of the engine is thereby reduced with a resulting decrease in aircraft speed and a further decrease in the ram pressure available. Two reasons are suggested why these circumstances do not occur with the inlet having slots and the modified internal duct shape: The slots permit the flow rate through the inlet to adjust itself to changes in pressure and thus damp fluctuations in the mass flow; and, since the internal shock losses for the usual operating condition occur in the portion of the duct where the change in area is small, the magnitude of the losses can change only slightly.\n\nCONCLUSIONS\n\nTests at Mach numbers between 1.36 and 2.01 of models having twin-scoop inlets situated on the sides of a long forebody indicated that maximum total-pressure ratios greater than those of a normal shock wave could be attained through the Mach number range and that pressure recovery within 2 percent of that associated with nose inlets could be attained at Mach numbers less than about 1.7. These relatively large total-pressure ratios resulted from an internal duct shape that improved the conditions for boundary-layer flow in the diffuser. The variation of total-pressure ratio with mass-flow ratio for the duct system indicated more stability than is usual with other types of inlets.\n\nAmes Aeronautical Laboratory,\nNational Advisory Committee for Aeronautics,\nMoffett Field, Calif.\n\nAPPENDIX\n\nSUBSONIC DIFFUSER WITH THE LOCAL PRESSURE GRADIENT\n\nPROPORTIONAL TO THE LOCAL STATIC PRESSURE\n\nIn the following analysis, it is assumed that the flow through the subsonic diffuser is unidimensional and that the relations for isentropic flow of a perfect gas are applicable.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:32.350711+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 23, "total_pages": 29, "image_filename": "19930085899_p23.jpg", "text": "22\nNACA RM No. L9A21\n\nFence\no Off\nΔ Small\n\nBending-moment coefficient, $C_B$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:42:37.758782+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 47, "total_pages": 72, "image_filename": "19930085491_p47.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:42:39.168845+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 19, "total_pages": 52, "image_filename": "19930085911_p19.jpg", "text": "18\nCONFIDENTIAL\nNACA RM E9F22\n\n$$\n\\frac{P_6}{P_5} = \\frac{\\left(1 + \\gamma_5 M_5^2\\right)}{\\left(1 + \\gamma_6 M_6^2\\right)} \\frac{\\left(1 + \\frac{\\gamma_6 - 1}{2} M_6^2\\right)^{\\frac{\\gamma_6}{\\gamma_6 - 1}}}{\\left(1 + \\frac{\\gamma_5 - 1}{2} M_5^2\\right)^{\\frac{\\gamma_5}{\\gamma_5 - 1}}} \\quad (15)\n$$\n\nWhen $M_7$ is less than 1.0 (no choking at outlet), $T_6$ is assumed and the corresponding $\\gamma_6$ is determined for the known fuel-air ratio. The quantity $\\left(1 + \\frac{W_f}{W_a}\\right)^2 \\frac{T_6}{T_5}$ is determined and $M_6$ and $P_6$ are determined from equations (14) and (15). The assumptions $P_6 = P_7$ and $T_6 = T_7$ are made.\n\n$$\n\\frac{A_{cr,7}}{A_7} = \\left(\\frac{A_{cr,6}}{A_6}\\right) \\left(\\frac{A_6}{A_7}\\right) \\quad (16)\n$$\n\nMach number $M_7$ can thus be determined from equations (8) and (16). Total pressure $P_6$ is the product of equations (12) and (15) and is assumed equal to $P_7$. It is then possible to determine the value of $p_7$ from equation (3). This value should equal the measured value of $p_7$ from the telemeter records. If they do not agree, the procedure is repeated, assuming a different value for $T_6$. When $M_7$ is equal to unity (choking at engine outlet), the total temperature after combustion is found directly from the relation for mass continuity\n\n$$\nT_7 = \\frac{g \\gamma_7 (\\gamma_7 + 1) A_7^2}{2 R_7} \\frac{p_7^2}{(W_a + W_f)^2} \\quad (17)\n$$\n\nAssumed values of $\\gamma_7$ and $R_7$ are used and then checked to agree with $T_7$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:42:39.357921+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 43, "total_pages": 149, "image_filename": "19930083192_p43.jpg", "text": "NACA TN 1976\n39\n\nof the O-2H airplane can be found in reference 34. Because of the use of improved instruments, the results from work with the XB-15 airplane are considered better than those obtained on the O-2H airplane.\n\nDiscussion\n\nWing area.- The results presented in figures 35 and 38(b) show good agreement between experiment and the detailed calculation based on net wing area. If gross area were used, the calculated values would be higher than the experimental values for a sharp gust. For a gradient gust the reverse would be true since the amount of alleviation due to pitch would be increased. A similar result applies to the canard airplane of figure 38(b) as indicated by table II of reference 15, which shows a decrease in acceleration of 30 percent for a gradient distance of 16 chords ($A = \\infty$).\n\nMass parameter.- Analytical studies that are based on extensive equations indicate that the mass parameter is the significant variable in that it affects the amount of alleviation due to vertical motion. The mass-parameter term has a negligible effect for a sharp-edge gust but is important for long gradient distances. This result is due to the fact that for a sharp gust, the airplane does not acquire much vertical velocity, but in a long gust the vertical velocity approaches the gust velocity.\n\nIn addition to its effect on the vertical motion of the airplane, an increase in the mass parameter for a given airplane, by increasing the gross weight, unexpectedly increases the alleviation of the total wing load due to pitch. (See fig. 36.) As the mass parameter increases, the resistance of the airplane to rotation would be expected to increase and therefore the amount of pitch and the alleviation due to the pitching motion would be decreased. The result just noted arises from the fact that the pitching moments on the airplane arise not only from the action of the gust but also from the vertical motion of the airplane. By reducing the vertical motion of the airplane, the adverse pitching moment is decreased, and thus the favorable effect predominates and provides more alleviation.\n\nIn some cases, such as in computing the factor K, wing loading instead of mass parameter has been taken as the significant variable. Such a substitution presumes a fixed relation between wing loading and mass parameter. Figure 39 indicates that this substitution may result in either underestimation or overestimation of imposed acceleration increments. The answer in any case depends on the relation between wing loading and mass parameter. Figure 39 shows that the substitution of wing loading appears satisfactory for transports but results in too-low acceleration increments for personal airplanes and too-high values for flying boats.", "timestamp": "2026-07-22T06:42:39.982977+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 29, "total_pages": 33, "image_filename": "19930085900_p29.jpg", "text": "```markdown\n28\nNACA RM L9D20\n\nSta. 10\nCONFIDENTIAL\n42\n\nSide views of model\n\nTrim, deg\nChine strips\nChine jets - 1/4-inch spacing\n\nEffective hydrodynamic lift and load on water, lb\nLoad on water\n\nResistance, lb\nNACA\n\nSpeed, fps\n\nFigure 13.- Comparison of strips and jets simulating chines; station 10\nto 42.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:42:40.836615+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 30, "total_pages": 92, "image_filename": "19930085870_p30.jpg", "text": "```markdown\nNACA RM No. L9D07\n31\n\nCONFIDENTIAL\n\n<!-- Image (62, 110, 914, 999) -->\n\n(a) Wing 1. w=0.223. R=1,390,000.\nFigure 5.- Aerodynamic characteristics of 8-\npercent-thick triangular wings at M=1.62.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:42:42.918499+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 15, "total_pages": 22, "image_filename": "19930085922_p15.jpg", "text": "14\nNACA RM No. L9C23\n\n<!-- Image (76, 110, 922, 901) -->\n\nFigure 5.- Variation with Mach number of the rolling-moment characteristics of the test wing for various angles of attack and aileron deflections; vertical fins off; $\\delta_{a_r} = 0$.", "timestamp": "2026-07-22T06:42:48.745231+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 2, "total_pages": 30, "image_filename": "19930085970_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:42:49.195776+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 78, "total_pages": 114, "image_filename": "19930086061_p78.jpg", "text": "74\nNACA RM L9J07\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\n(a) $\\psi = 0^\\circ$.\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\n10°\n\n(b) $\\psi = 10^\\circ$.\n\nNACA\n\nFigure 26.- Pressure distribution about wing 3 at various angles of yaw;\n$\\alpha = 50.1^\\circ$.", "timestamp": "2026-07-22T06:42:52.658089+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 53, "total_pages": 62, "image_filename": "19930082918_p53.jpg", "text": "64\nNACA TN 1940\n\nBrinell hardness\nAging time, hr\n\n[Figure: Graph showing Brinell hardness vs. Aging time, hr. Three curves are plotted:\n- Aged at 1600° F\n- Aged at 1400° F\n- Aged at 1200° F\nThe x-axis ranges from 0 to 1000 hr. The y-axis ranges from 180 to 260 Brinell hardness. Data points are marked with circles, squares, and crosses. The NACA logo is in the top right corner of the graph.]\n\nFigure 13.- Effect of aging on hardness of low-carbon N-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:42:54.191402+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 1, "total_pages": 30, "image_filename": "19930085975_p1.jpg", "text": "```markdown\nFILE COPY\nNO 5\n\nCONFIDENTIAL\n\nCopy 392\nRM L9E10\n\nNACA RM L9E10\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEFFECTS OF MACH NUMBER AND SWEEP ON THE DAMPING-IN-ROLL\nCHARACTERISTICS OF WINGS OF ASPECT RATIO 4\n\nBy\n\nRichard E. Kuhn and Boyd C. Myers, II\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THIS OFFICE AS SOON AS POSSIBLE.\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n1512 H STREET, N. W.\nWASHINGTON 25, D. C.\n\nCLASSIFICATION CHANGED TO\nCLASSIFIED DOCUMENT\n\nThis document contains information affecting the national defense of the United States within the meaning of the Espionage Act, 50 U.S.C. 31 and 32, as amended. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law.\n\nDISTRIBUTION\nonly 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 having security clearance who of necessity must be informed thereof.\n\nDATE 8-13-54\n\nW. CHOWLEY\n\nE.L.B.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nJune 27, 1949\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:42:56.971954+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 70, "total_pages": 78, "image_filename": "19930082483_p70.jpg", "text": "```markdown\n68\n\n500\n400\n300\n200\n100\n0\n\nCorrected turbine output, hp\n\nOver-all turbine\nefficiency, $\\eta'$ .60\n.65\n.70\n.73\n.75\n.75\n.75\n.73\n.70\n.65\n.60\n.50\n.40\n\nCorrected rotor speed, $N/\\sqrt{\\theta_1}$, rpm\n3500\n4500\n5500\n6500\n7500\n8500\n9500\n10,500\n11,500\n12,500\n\nTotal-pressure ratio, $p_1/p_{2e}$\n1.5\n1.6\n1.8\n2.0\n2.2\n2.4\n\n500\n400\n300\n200\n100\n0\n\n(a) Full admission.\n\nNACA\n\nFigure 8. - Performance characteristics of gas turbine. Corrected inlet total pressure, 29.92 inches mercury\nabsolute; corrected inlet total temperature, 518.6° R; turbine rotor pitch-line diameter, 1.166 feet; flow area\nat turbine-inlet measuring station, 1.60 square feet; flow area at turbine-discharge measuring station (for full\nadmission), 0.686 square foot. Turbine design operating point, D.\n\nNACA TN NO. 1807\n\n1032\n```", "timestamp": "2026-07-22T06:42:57.162432+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 80, "total_pages": 98, "image_filename": "19930086073_p80.jpg", "text": "78\nNACA RM A9H04\n\n<!-- Image (123, 116, 898, 797) -->\n\n$$\n\\begin{array}{cccc}\n\\bigcirc & \\square & \\diamond & \\triangle \\\\\n0.0 & 6.0 & 12.0 & 15.9\n\\end{array}\n$$\nAngle of sideslip, $\\beta$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 17. — Continued.", "timestamp": "2026-07-22T06:42:57.792693+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 43, "total_pages": 60, "image_filename": "19930085862_p43.jpg", "text": "NACA RM No. L9A07\n41\n\n<!-- Image (162, 119, 885, 909) -->\n\nFigure 12.- Variation of aileron hinge-moment coefficient with aileron deflection for various flap configurations.", "timestamp": "2026-07-22T06:42:59.209657+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 43, "total_pages": 46, "image_filename": "19930085548_p43.jpg", "text": "42\nNACA RM No. E8L30\n\nHeat loss to coolant\n(percent of heat input)\n\nExhaust-gas temperature, °F\n\nInlet-manifold\npressure\n(lb/sq in. abs.)\nO 80\n□ 100\n◇ 120\n△ 135\n\nFuel-air ratio\n\nFigure 20. - Variation of heat loss and exhaust-gas\ntemperature of experimental cylinder with fuel-air\nratio. Compression ratio, 5.25; inlet-manifold\ntemperature, 400° F.\n\n[Figure: A graph with two y-axes. The left y-axis is labeled \"Heat loss to coolant (percent of heat input)\" ranging from 20 to 50. The right y-axis is labeled \"Exhaust-gas temperature, °F\" ranging from 400 to 2000. The x-axis is labeled \"Fuel-air ratio\" ranging from 0 to .05. The graph contains multiple data series represented by circles, squares, diamonds, and triangles, with trend lines drawn through them. A legend inside the graph indicates the symbols correspond to different inlet-manifold pressures. A NACA logo is present in the bottom right corner of the plot area.]", "timestamp": "2026-07-22T06:42:59.623696+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 2, "total_pages": 55, "image_filename": "19930085966_p2.jpg", "text": "NACA RM L9B17 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL DETERMINATION OF THE SUBSONIC PERFORMANCE\nOF A RAM-JET UNIT CONTAINING THIN-PLATE BURNERS\nBy John R. Henry\n\nSUMMARY\n\nThe performance of a ram-jet unit consisting of an intake diffuser, an exhaust nozzle, and a cluster of thin-plate burners contained in a semicircular combustion chamber was investigated in the Langley induction aerodynamics laboratory. Data were taken over a fuel-air-ratio range from 0 to 0.049, a fuel flow range from 0 to 3100 pounds per hour, at combustion-chamber inlet velocities from 40 to 195 feet per second, and at simulated free-stream Mach numbers from 0.20 to 0.55.\n\nCombustion efficiencies from 56 to 72 percent were obtained. At the higher fuel flows investigated, marked decreases in combustion efficiency resulted from increases in fuel flow. This characteristic led to the conclusion that operation under high-thrust-output conditions would not be feasible. It was estimated that the combustion-chamber performance obtained in the subsonic test-stand investigation would produce at supersonic flight speeds thrust coefficients regarded as too low to be practical.\n\nThe cycle-efficiency and propulsive-efficiency product of the ram-jet unit was approximately 80 percent of that for a no-pressure-loss unit under the same conditions of operation.\n\nThe performance of the intake diffuser, which had an area ratio of 2.14 to 1 and an equivalent conical angle of expansion of $16^\\circ$, was a unique function of inlet-boundary-layer thickness. Over 99 percent diffuser efficiency was obtained when the boundary layer at the inlet was completely eliminated.\n\nINTRODUCTION\n\nThis paper is concerned with the determination of subsonic performance characteristics of the ram-jet burner and combustion-chamber assembly shown in figure 1. The burners and semicircular combustion chamber were designed in 1942 for application as a speed booster to be mounted on the under side\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:04.537205+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 9, "total_pages": 92, "image_filename": "19930085951_p9.jpg", "text": "NACA RM L9D29\nCONFIDENTIAL\n7\n\nthis group, the values of induced efficiency may be considered accurate to within about 1 percent for all three propellers. At the design value of advance ratio (2.1) the induced losses amount to about 4 percent, and the profile-drag losses amount to 3 percent for the NACA 10-(3)(08)-03 propeller. At this same design value of advance ratio the profile-drag losses of the NACA 10-(3)(08)-03R and NACA 10-(3)(12)-03 propellers are about twice as great as those of the propeller having the thinner blade sections. The higher efficiency of the NACA 10-(3)(08)-03 propeller reflects the importance of using thin, efficient airfoil sections throughout the blade.\n\nA comparison of the envelope efficiencies of the NACA propellers in the 0.045-solidity group is shown in figure 27 for various rotational speeds. Again the thinnest blade (NACA 10-(3)(05)-045) of the group has the highest efficiency, and the envelope efficiencies of all the propellers are reduced at the higher rotational speeds. There are only small differences in the envelope efficiencies of the NACA 10-(3)(05)-045 and NACA 10-(3)(062)-045A propellers, although the blade sections of the latter propeller are thicker at all radii except at the shank and at the tip. The fact that the thinner blade sections of the NACA 10-(3)(05)-045 propeller do not improve its efficiency much above that of the NACA 10-(3)(062)-045A propeller may possibly be explained by only slight improvements in the lift-drag ratios of the thinner blade sections. Figure 14 of reference 11 shows that there is little difference in the lift-drag ratios of 6- and 9-percent-thick sections (16-3xx airfoils) at lift coefficients up to 0.4. Since a large portion of the radial load is perhaps carried by the blade sections of the thinner NACA propellers having a thickness of 9 percent or less, a reduction in thickness from 6.2 to 5 percent at the 0.7 radius might be expected to cause only small changes in the propeller efficiency. This is perhaps true for the conditions of operation under which the tests were made; however, it should be pointed out that the differences in efficiency between the two propellers may be greater at air-stream Mach numbers higher than those used in these tests. If higher air-stream Mach numbers and lower rotational speeds had been used to attain the helical-tip Mach numbers shown in the figures, then greater portions of the blades would be subjected to the effects of compressibility. Since the airfoil data in reference 11 show that, in general, the thinner sections have the higher lift-drag ratios at the higher Mach numbers, then a reasonable assumption would be that the propeller having the thinner blade sections along the radius should have less efficiency losses due to compressibility. The envelope efficiencies of the NACA 10-(3)(062)-045 and NACA 10-(3)(08)-045 propellers, which had the thickest blade sections, are from $1\\frac{1}{2}$ to 4 percent lower than the envelope efficiencies of the thinnest propeller in the 0.045-solidity group.\n\nThe optimum, or induced, efficiency has been calculated using the same values of power coefficients as were obtained for each of the NACA propellers in the 0.045-solidity group, and the curve showing induced efficiency in figure 27(b) may be considered accurate to within about 1 percent for all four propellers. At the design value of advance\n\nUNCLASSIFIED\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:04.985135+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 24, "total_pages": 29, "image_filename": "19930085899_p24.jpg", "text": "NACA RM No. L9A21\n\n$\\frac{q_{\\text{wake}}}{q}$\n\nM = 0.70 $\\alpha = 10^\\circ$\n\nM = 0.80 $\\alpha = 10^\\circ$\n\nM = 0.90\n\nWing alone\n\nWing-fuselage\n\n$\\frac{q_{\\text{wake}}}{q}$\n\n$\\alpha = 4^\\circ$\n\n$\\alpha = 4^\\circ$\n\n$\\frac{q_{\\text{wake}}}{q}$\n\n$\\alpha = 0^\\circ$\n\n$\\alpha = 0^\\circ$\n\n-80 -40 0 40 80 -80 -40 0 40 80 -80 -40 0 40 80\n\nTail height, $h_t$, percent semispan\n\nNACA\n\nFigure 10.- Dynamic-pressure surveys in region of tail plane for a model with $45^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n23", "timestamp": "2026-07-22T06:43:05.352180+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 11, "total_pages": 22, "image_filename": "19930085928_p11.jpg", "text": "10 CONFIDENTIAL NACA RM No. A9A31\n\nThe ratio of static to total pressure in terms of the local Mach number is indicated by the equation\n\n$$\n\\frac{p}{H} = \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right)^{-\\frac{\\gamma}{\\gamma - 1}} \\tag{A1}\n$$\n\nDifferentiating this expression with respect to $x/L$ and collecting terms gives the following equation for the local static-pressure gradient\n\n$$\n\\frac{dp}{d(x/L)} = - H \\gamma \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right)^{\\frac{-2\\gamma + 1}{\\gamma - 1}} M \\frac{dM}{d(x/L)} \\tag{A2}\n$$\n\nThe ratio of the pressure gradient to the local static pressure is then\n\n$$\n\\frac{dp/[d(x/L)]}{p} = - \\frac{\\gamma H}{p} \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right)^{\\frac{-2\\gamma + 1}{\\gamma - 1}} M \\frac{dM}{d(x/L)} \\tag{A3}\n$$\n\nTaking $\\frac{dp/[d(x/L)]}{p} = K$, a constant for a diffuser of given length, and substituting for $\\frac{H}{p}$ in equation (A3) yields\n\n$$\nK = - \\gamma \\left(1 + \\frac{\\gamma - 1}{2} M^2\\right)^{-1} M \\frac{dM}{d(x/L)} \\tag{A4}\n$$\n\nAssuming that, when $x/L = 0$, $M = 1.0$, integration of equation (A4) gives\n\n$$\nK \\left(\\frac{x}{L}\\right) = - \\frac{\\gamma}{\\gamma - 1} \\ln \\left[ \\left(\\frac{2}{\\gamma - 1} + M^2\\right) \\left(\\frac{\\gamma - 1}{\\gamma + 1}\\right) \\right] \\tag{A5}\n$$\n\nCoefficient $K$ is evaluated from equation (A5) by selecting a value for $M$ at $x/L = 1$. This value of $M$ is usually determined by the permissible settling-chamber velocity.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:05.535924+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 48, "total_pages": 72, "image_filename": "19930085491_p48.jpg", "text": "NACA RM No. A8J04\n\nCONFIDENTIAL\n\n[Figure: 3D diagram of lift distribution on a tapered, flat plate with leading edge swept within the Mach cone. Arrows labeled “α” and “-α” indicate angle of attack directions. Shaded regions represent pressure or lift distribution across the plate surface.]\n\nNACA  \nA-13463\n\nFigure 5.— Lift distribution on tapered, flat plate with leading edge swept within the Mach cone.\n\n47\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:07.465353+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 31, "total_pages": 92, "image_filename": "19930085870_p31.jpg", "text": "```markdown\n32\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\nElliptical L.E. { O CL\n                 { □ Cm\nWedge L.E.      { △ CL\n                 { ◇ Cm\n\n.24\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n\n.01\nCm\n0\n-.01\n\nElliptical L.E. { O CD\n                 { □ L/D\nWedge L.E.      { △ CD\n                 { ◇ L/D\n\n.06\n.04\nCD\n.02\n0\n\n6\n4\nL/D\n2\n0\n\n-8  -6  -4  -2   0   2   4   6   8\n          α, deg\n\n(b) Wing 2. w = 0.412; R = 1,390,000.\nFigure 5. - Continued.\nCONFIDENTIAL\n\n```", "timestamp": "2026-07-22T06:43:15.839542+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 81, "total_pages": 98, "image_filename": "19930086073_p81.jpg", "text": "NACA RM A9H04\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 17. – Concluded.\n\nNACA\n\n79", "timestamp": "2026-07-22T06:43:19.898768+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 71, "total_pages": 78, "image_filename": "19930082483_p71.jpg", "text": "```markdown\nNACA TN NO. 1807\n\nCorrected turbine output, hp\n\nOver-all turbine efficiency, $\\eta'$ .60\nFull admission\n\nCorrected rotor speed, $N/\\sqrt{\\theta_1}$, rpm\n\nTotal-pressure ratio, $p_1/p_2$\n\nOver-all turbine efficiency, $\\eta'$\n120° admission\n\nCorrected rotor speed, $N/\\sqrt{\\theta_1}$, rpm\n\nTotal-pressure ratio, $p_1/p_2$\n\n(b) Full and 120° admission.\n\nNACA\n\nFigure 8. - Concluded. Performance characteristics of gas turbine. Corrected inlet total pressure, 29.92 inches mercury absolute; corrected inlet total temperature, 518.6° R; turbine rotor pitch-line diameter, 1.166 feet; flow area at turbine-inlet measuring station, 1.60 square feet; flow area at turbine-discharge measuring station (for full admission), 0.686 square foot. Turbine design operating point, D.\n\n69\n```", "timestamp": "2026-07-22T06:43:21.204624+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 44, "total_pages": 46, "image_filename": "19930085548_p44.jpg", "text": "NACA RM No. E8L30\n43\n\nExhaust-gas temperature, °F\nFuel-air ratio\n\nInlet-manifold pressure\n(lb/sq in. abs.)\n135\n120\n100\n80\nExperimental\ndata\n--- Calculated data,\nequation (5)\n\n[Figure: Graph showing exhaust-gas temperature vs. fuel-air ratio with experimental data points and a calculated data line.]\n\nFigure 21. - Comparison of experimental and calculated exhaust-gas temperatures. Inlet-manifold temperature, 400° F; compression ratio, 5.25. Calculation made at inlet-manifold pressure of 100 pounds per square inch.", "timestamp": "2026-07-22T06:43:27.048846+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 30, "total_pages": 33, "image_filename": "19930085900_p30.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 29\n\n[Figure: (a) 1/4-inch-spaced jets; trim, 6.3°.]\n\n[Figure: (b) Strips; trim, 6.6°.]\n\nFigure 14.- Comparison of jets and strips at 35 feet per second; simulated chines; station 10 to 42.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:27.257516+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 20, "total_pages": 52, "image_filename": "19930085911_p20.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 19\n\nCombustion efficiency is determined by the equation\n\n$$\n\\eta_b = \\frac{\\left(1 + \\frac{W_F}{W_a}\\right) H_g - H_a}{\\frac{W_F}{W_a} h + \\frac{Q}{W_a}}\n$$\n\n(18)\n\nThe enthalpy values are obtained from reference 5.\n\nThrust coefficient is defined as\n\n$$\nC_F = \\frac{2F_n}{\\gamma_0 p_0 M_0^2 A_{\\text{max}}}\n$$\n\n(19)\n\nwhere\n\n$$\nF_n = \\left(\\frac{W_a + W_F}{g}\\right) V_7 - \\left(\\frac{W_a}{g}\\right) V_0 + A_7 (p_7 - p_0)\n$$\n\n(20)\n\n$$\nV_7 = M_7 \\left( \\frac{\\gamma_7 g R_7 T_7}{1 + \\frac{\\gamma_7 - 1}{2} M_7^2} \\right)^{\\frac{1}{2}}\n$$\n\n(21)\n\nThe weight of the ram jet at any time during the flight is\n\n$$\nW = W_i - \\int_0^T W_F \\, d\\tau\n$$\n\n(22)\n\nThe external drag is determined as\n\n$$\nD = F_n - W a_n\n$$\n\n(23)\n\nThe external drag coefficient is defined as\n\n$$\nC_D = \\frac{D}{\\frac{\\gamma_0}{2} p_0 M_0^2 A_{\\text{max}}}\n$$\n\n(24)\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:27.257644+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 14, "total_pages": 34, "image_filename": "19930085913_p14.jpg", "text": "```markdown\nTABLE I.- EXPERIMENTAL DATA\n\n| Model | Run | $q_r$ (lb/sq ft) | $V_r$ (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** | | | | **2nd natural mode** | | | | **3rd natural mode** | | | | **Flutter mode** | | | |\n| | | | | | | 1st | 2nd | 3rd | Flutter | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) |\n| Model A<br>Unswept untapered wing<br>Weight moved along leading<br>edge; $\\theta_w = -1$<br>Reynolds number $\\approx 3653.1 V_r$ | 1 | 33.29 | 175.8 | 0.1515 | 0 | 2.24 | 13.65 | 19.64 | 21.00<br>$\\theta_{12.00}$ | --- | Ref | - | 180 | Ref | --- | 0 | --- | --- | -- | -- | --- |\n| | 2 | 16.05 | 121.5 | .1045 | 0 | 2.24 | 13.65 | 19.64 | 14.75 | --- | Ref | - | 180 | Ref | --- | 0 | --- | Ref | 36 | 0 | 216 |\n| | 3 | 16.99 | 125.1 | .1075 | 5.55 | 2.26 | 13.64 | 19.38 | 14.80 | --- | Ref | - | 180 | Ref | --- | 0 | --- | Ref | 35 | 0 | 217 |\n| | 4 | 28.21 | 161.5 | .1389 | 11.11 | 2.26 | 12.88 | 17.48 | 10.04 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Ref | 0 | 0 | 0 |\n| | 5 | 27.86 | 160.5 | .1379 | 16.66 | 2.25 | 10.78 | 16.24 | 8.51 | Ref | --- | - | --- | Ref | 0 | 0 | 180 | Ref | 32 | 0 | 32 |\n| | 6 | 28.30 | 161.9 | .1390 | 22.20 | 2.23 | 8.83 | 16.15 | 7.41 | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | Ref | 50 | 22 | 356 |\n| | 7 | 33.91 | 177.6 | .1525 | 27.77 | 2.20 | 7.82 | 16.25 | 6.51 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Ref | 36 | 35 | 3 |\n| | 8 | 35.18 | 172.3 | .1480 | 33.33 | 2.14 | 7.11 | 16.25 | Divergence | Ref | 0 | - | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 38.90 | 2.07 | 6.85 | 15.98 | Divergence | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 44.40 | 1.97 | 6.82 | 15.75 | Divergence | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 50.00 | 1.85 | 7.11 | 15.46 | Divergence | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 55.50 | 1.74 | 7.52 | 15.01 | Divergence | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 61.11 | 1.63 | 8.18 | 14.62 | Divergence | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 66.66 | 1.53 | 8.82 | 14.22 | Divergence | Ref | --- | 0 | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | -- | ----- | ----- | ----- | 72.20 | 1.44 | 9.61 | 13.80 | Divergence | Ref | - | 0 | 0 | Ref | 0 | 0 | 180 | --- | -- | -- | --- |\n| | 9 | 35.49 | 181.9 | .1560 | 80.50 | 1.38 | 10.08 | 13.62 | Divergence | Ref | 0 | 0 | 0 | Ref | 0 | - | 180 | --- | -- | -- | --- |\n| | 10 | 22.94 | 145.8 | .1250 | 83.30 | 1.27 | 10.12 | 13.68 | 11.80 | Ref | 0 | 0 | 0 | Ref | 180 | 0 | 0 | Ref | 99 | 0 | 319 |\n| | 11 | 12.34 | 106.9 | .0915 | 88.90 | 1.18 | 9.77 | 13.85 | $^a$12.09 | Ref | 0 | 0 | 180 | Ref | 180 | 0 | 0 | Ref | 43 | 0 | --- |\n| | 12 | 10.62 | 98.6 | .0845 | 94.40 | 1.12 | 9.21 | 14.01 | 12.09 | Ref | 0 | 0 | 180 | Ref | 180 | 0 | 0 | Ref | 123 | 0 | 230 |\n| | 13 | 10.28 | 97.4 | .0835 | 98.60 | 1.06 | 8.69 | 14.03 | 12.12 | Ref | 0 | 0 | 180 | Ref | 180 | 0 | 0 | Ref | 77 | 0 | 337 |\n\n[Figure: Diagram of wing model showing length l, thickness t, and distance 2b]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\nNACA RM 1924\n13\n```", "timestamp": "2026-07-22T06:43:28.351132+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 16, "total_pages": 22, "image_filename": "19930085922_p16.jpg", "text": "NACA RM No. L9G23\n\nRolling-moment coefficient, $C_l$\n\nMach number, $M$\n\n$\\delta_{a_R}$ (deg)\n□ 44\n∇ 18\n○ 0\n∇ -18\n△ -44\n\n[Figure: Graph showing variation of rolling-moment coefficient with Mach number for different aileron deflections. Data points are plotted for δaR = 44°, 18°, 0°, -18°, and -44°. NACA logo appears near bottom right of plot area.]\n\nFigure 6.- Variation with Mach number of the rolling-moment characteristics of the test wing, α = 0.30°; vertical fins on; δeT = 0.\n\n15", "timestamp": "2026-07-22T06:43:28.796712+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 44, "total_pages": 60, "image_filename": "19930085862_p44.jpg", "text": "42\nNACA RM No. L9A07\n\n.16\n.12\n.08\n.04\n0\n$C_{h\\alpha}$ -.04\n-.08\n-.12\n-.16\n-.20\n\n(d) Leading-edge\nand split flaps.\n\n(deg)\nO 0.4\n$\\square$ 4.8\n$\\diamond$ 9.0\n$\\Delta$ 13.2\n\n.16\n.12\n.08\n.04\n0 $C_{h\\alpha}$\n-.04\n-.08\n-.12\n-.16\n-.20\n\n(e) Leading-edge and\nsplit flaps and fences.\n\n(deg)\nO 0.6\n$\\square$ 4.6\n$\\diamond$ 9.0\n$\\Delta$ 13.2\n\n.16\n.12\n.08\n.04\n0\n$C_{h\\alpha}$ -.04\n-.08\n-.12\n-.16\n-.20\n\n(deg)\nO 0.1\n$\\square$ 4.3\n$\\diamond$ 8.6\n$\\Delta$ 12.8\n\n(f) Leading-edge\nflaps and fences\n\n-25 -20 -15 -10 -5 0 5 10 15 20 25\n$\\delta_o$, deg\n\n[Figure: NACA logo]\n\nFigure 12.- Concluded.", "timestamp": "2026-07-22T06:43:28.997882+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 54, "total_pages": 62, "image_filename": "19930082918_p54.jpg", "text": "NACA TN 1940\n65\n\n[Figure: Graph showing creep rate vs. aging time for M-155 alloy under various aging temperatures]\n\nCreep rate, in./in./hr\nAging time, hr\n\nFigure 14.- Effect of aging on secondary creep rates at 30,000 psi and 1200° F of low-carbon M-155 alloy solution-treated 10 hours at 2200° F and water-quenched.", "timestamp": "2026-07-22T06:43:38.264227+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 10, "total_pages": 92, "image_filename": "19930085951_p10.jpg", "text": "8\nUNCLASSIFIED\nCONFIDENTIAL\nNACA RM L9D29\n\nratio (2.1) the curves in figure 27(b) show that the induced losses amount\nto about 5 percent, and the profile-drag losses amount to only 2 percent\nfor the NACA propeller having the thinnest blade sections. The\nNACA 10-(3)(08)-045 propeller, which has the thickest outboard blade\nsections in this group, suffers the greatest profile-drag loss (about\n4 percent).\n\nThe envelope efficiencies of all the NACA propellers in both solidity\ngroups are quite high, and the differences in efficiency between the various\npropellers of each group are small and difficult to analyze for some condi-\ntions of operation. Where the differences in efficiency are small the\nrelative differences in thrust and power coefficients are also small, and\nit is difficult to draw any general conclusion as to whether a loss in\nefficiency is caused by a loss of thrust, or an increase of power, or both.\n\nEffect of thickness ratio on constant-power propeller operation.-\nAirplane propellers often operate over an extensive range of advance ratio\nat constant rotational speed and torque. Since blade-section thickness\nratio affects the power-absorption qualities of a propeller to some extent,\nthe data for the NACA propellers have been compared in figure 28 for\noperation at a constant power coefficient of 0.15 and a rotational speed\nof 1140 rpm. This condition of operation may be considered representative\nfor two-blade, 10-foot-diameter propellers and provides a reasonable basis\nof comparison. It is interesting to note in figure 28(a) that the drop in\nefficiency occurs at a lower value of advance ratio but is more gradual\nfor the round-shank propeller than for the NACA 10-(3)(12)-03 propeller,\nwhich has thinner shanks but thicker outboard blade sections than the\nround-shank propeller. At advance ratios above 3.2 it appears that the\nthicker outboard sections of the NACA 10-(3)(12)-03 propeller may cause\ngreater efficiency losses than the thick shank sections of the\nNACA 10-(3)(08)-03R propeller. However, the gain in efficiency of about\n10 percent at an advance ratio of 3.2 ($M = 0.54$), which may be realized\nby using the thinner blade sections of the NACA 10-(3)(08)-03 propeller,\nshould be emphasized. A reduction in blade-section thickness from 12 to\n8 percent at the 0.7 radius, or approximately one-third all along the\nradius, resulted in gains in efficiency up to 10 percent. A reduction in\nblade-section thickness of only the inboard blade sections from 30 to\n13 percent at the 0.3 radius also resulted in gains in propeller efficiency\nup to 10 percent.\n\nFigure 28(b) compares the efficiencies for constant power operation\nof the propellers in the 0.045-solidity group. The differences in effi-\nciency of the propellers in this group do not amount to more than 4 percent,\nand their efficiencies appear to be about equal at the lowest values of\nadvance ratio and also at the highest values of advance ratio. The single\nexception is the propeller which has the thickest shank sections\n(NACA 10-(3)(062)-045); at the highest value of advance ratio (3.5) the\nefficiency of this propeller is about 2 percent less than the efficiency\nof the other three propellers. The curves in figure 28(b) show that a\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:39.712606+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 3, "total_pages": 55, "image_filename": "19930085966_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9B17\n\nof a fighter airplane. Initial tests were run in the latter part of 1942\nin a 3-foot combustion wind tunnel at the Langley Laboratory. To adapt\nthe model to the tunnel a nozzle was placed upstream of the combustion\nchamber, and to obtain the maximum air flow for the power available the\nproducts of combustion were discharged through a diffuser. Due to low\ntunnel power and lack of instrumentation not many significant quantitative\nresults were obtained; however, crude measurements indicated a 50-percent\ncombustion efficiency at a fuel-air ratio of 0.025 with an inlet velocity\nof 75 feet per second.\n\nAlthough rocket developments soon outmoded the speed-booster appli-\ncation of the thin-plate burner, the performance under high-thrust-output\nconditions was of interest for possible application to supersonic aircraft.\nWhen the blower facilities of the Langley induction aerodynamics laboratory\nbecame available in 1945, an investigation was initiated to obtain more\ncomprehensive burner performance using a test setup simulating as closely\nas possible a flight configuration. The simulation consisted of replacing\nthe intake nozzle with an intake diffuser and the exhaust diffuser with an\nexhaust nozzle and bleeding off the boundary layer at the diffuser inlet.\nPreliminary tests were run in which the burners were modified to obtain\napproximately the maximum performance for the present burner configuration.\nThe use of two 1000-horsepower centrifugal blowers and a high-capacity,\npositive-displacement fuel pump permitted testing over a wide range of\nfuel and air flows up to back pressures at the diffuser inlet corresponding\nto a simulated flight Mach number of 0.55. The data have been analyzed\nin a manner similar to that of reference 1. An estimate, based on the\nsubsonic test-stand data, of thrust coefficients at supersonic flight\nspeeds is presented.\n\nSYMBOLS\n\nThe following symbols are used throughout the paper:\n\n| | |\n| :--- | :--- |\n| A | cross-sectional area, square feet |\n| $C_F$ | thrust coefficient |\n| $c_p$ | specific heat at constant pressure, British thermal units per pound per degree Fahrenheit |\n| F | thrust, pounds |\n| g | acceleration due to gravity, feet per second per second |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:43:43.490715+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 44, "total_pages": 149, "image_filename": "19930083192_p44.jpg", "text": "40\nNACA TN 1976\n\nPhase lag.- The data from the general analysis were utilized to obtain the lag in chords of the acceleration peak behind the flat-top-gust peak. Figure 41 shows that the lag is primarily a function of the gradient distance for a given mass parameter although some variations due to the tail area and tail length are noted. No effect due to a variation in the center-of-gravity position from 15 to 35 percent of the chord is noted. The results indicate that for a mass parameter of approximately 10, the lag in the gust peak could be as much as 2 chords for an 8- to 10-chord gust. Computations indicate that this lag increases to 3 chords for a gradient distance of 8 chords when the mass parameter is about 30.\n\nIn the calculations for the triangular and sinusoidal gusts, the lag of the acceleration peak behind the gust peak for infinite mass parameter, the most adverse case, does not exceed 2 chords. The effect of the phase lag of the acceleration behind the gust can probably be neglected since, for many calculations of acceleration increments and load factors, a change of 1 chord in 8 has a negligible effect on the acceleration ratio.\n\nStatic margin.- Figure 34 indicates that the total wing load on a conventional airplane decreases with increasing stability and increasing gradient distance because of the resulting increase in pitch. The figure also shows that the method of obtaining a given static margin is more important than the actual value of the static margin. The results for tailless airplanes are discussed in the following section.\n\nAlthough no data are available for canard airplanes, a comparison of figures 38(b) and 32 (both airplanes have the same static margin) show a marked effect on the acceleration increment due to the radical change in configuration. From the results obtained, it is concluded that the effect of airplane pitch on the wing load cannot be specified by static margin alone.\n\nCenter-of-gravity position.- Inspection of figure 34 indicates that, for a given configuration, the alleviation due to pitch varies almost directly with the center-of-gravity position. The rate of change of pitch effect with center-of-gravity position is roughly independent of the configuration for the conditions analyzed and depends mainly on the gust-gradient distance. For a gust with a gradient distance of 8 chords, figure 34 indicates that a 1-percent change in center-of-gravity position gives a 1-percent change in load factor. The change in load is doubled for the gust with a gradient distance of 16 chords and about halved for a sharp-edge gust.\n\nThe results shown in figure 34, which are for a single value of the mass parameter, and those presented in figure 36 indicate that the effect of pitch on the gust load factor will increase as the mass parameter is", "timestamp": "2026-07-22T06:43:46.950173+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 45, "total_pages": 60, "image_filename": "19930085862_p45.jpg", "text": "```markdown\nNACA RM No. L9A07\n43\n\n<!-- Image (143, 109, 875, 872) -->\n\nFigure 13.- Effects of high-lift and stall-control devices on aileron effectiveness parameter $C_{l\\delta}$.\n```", "timestamp": "2026-07-22T06:43:53.688250+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 82, "total_pages": 98, "image_filename": "19930086073_p82.jpg", "text": "```markdown\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n\nLift coefficient, $C_L$\n\n0 0 0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\circ$ $\\square$ $\\diamond$ $\\triangle$\n0.0 6.0 12.0 15.9\nAngle of sideslip, $\\beta$, deg\n\n(a) $C_L$ vs $\\alpha$.\n\nFigure 18.— Wing plus body plus vertical tail at various angles of sideslip with flaps undeflected and rudder deflected $10^\\circ$.\n\n[NACA logo]\n\n80\nNACA RM A9H04\n```", "timestamp": "2026-07-22T06:44:12.060046+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 31, "total_pages": 33, "image_filename": "19930085900_p31.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:44:12.719101+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 32, "total_pages": 92, "image_filename": "19930085870_p32.jpg", "text": "NACA RM No. L9D07\n33\n\nCONFIDENTIAL\n\nElliptical L.E. {○ $C_L$\n                {□ $C_m$\nWedge L.E. {△ $C_L$\n           {◇ $C_m$\n\n$C_L$\n.24\n.16\n.08\n0\n-.08\n-.16\n-.24\n\n$C_m$\n.01\n0\n-.01\n\nElliptical L.E. {○ $C_D$\n                {□ $L/D$\nWedge L.E. {△ $C_D$\n           {◇ $L/D$\n\n$C_D$\n.06\n.04\n.02\n0\n\n$L/D$\n6\n4\n2\n0\n\n$\\alpha$, deg\n-8 -6 -4 -2 0 2 4 6 8\n\n(c) Wing 3. w=0.514; R = 1,380,000.\nFigure 5. - Continued.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:12.909652+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 72, "total_pages": 78, "image_filename": "19930082483_p72.jpg", "text": "70\nNACA TN No. 1807\n\nAdmission\n(deg)\n360\n270\n240\n180\n120\n90\n\nRotor-tip leakage loss corrected to sea level, hp\n20.0\n15.0\n10.0\n9.0\n8.0\n7.0\n6.0\n5.0\n4.0\n3.0\n2.5\n2.0\n1.5\n1.0\n\n(a) Rotor-tip leakage loss.\n\nBearing loss, hp\n15.0\n10.0\n9.0\n8.0\n7.0\n6.0\n5.0\n4.0\n3.0\n2.5\n2.0\n1.5\n1.0\n\nAll admissions\nAll pressure ratios\n\n(b) Bearing loss.\n\nAdmission\n(deg)\n90\n120\n180\n240\n270\n\nPumping loss corrected to sea level, hp\n10.0\n9.0\n8.0\n7.0\n6.0\n5.0\n4.0\n3.0\n2.5\n2.0\n1.5\n1.0\n\n(c) Pumping loss.\n\nDriving-fluid losses corrected to sea level, hp\n50\n40\n30\n25\n20\n15\n10\n9\n8\n7\n6\n5\n4\n3\n2\n\nCorrected rotor speed, $N/\\sqrt{\\theta_1}$, rpm\n2 3 4 5 6 7 8 9 10 12.5x10³\n\n(d) Driving-fluid losses.\n\nNACA\n\nFigure 9. - Calculated and experimental losses for gas turbine operating at total-pressure ratio of 2.0 for full and partial admissions.", "timestamp": "2026-07-22T06:44:16.603400+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 45, "total_pages": 46, "image_filename": "19930085548_p45.jpg", "text": "44\nNACA RM No. E8L30\n\nExhaust-gas temperature, $^\\circ$R\n\nCompression ratio\n7.00\n\nExperimental\nCalculated\n--- (equation (5))\n\n5.25\n\n4.50\n\n4.00\n\n[Figure: NACA logo]\n\nFuel-air ratio\n\nFigure 22. - Comparison of experimental and calculated values of exhaust-gas temperature at various compression ratios and fuel-air ratios. Inlet-manifold temperature, 400$^\\circ$ F; inlet-manifold pressure, 100 pounds per square inch absolute.", "timestamp": "2026-07-22T06:44:17.743057+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 4, "total_pages": 55, "image_filename": "19930085966_p4.jpg", "text": "NACA RM L9B17 CONFIDENTIAL 3\n\n| Symbol | Definition |\n| :--- | :--- |\n| $h_c$ | lower heating value of fuel (19,000 Btu/lb) |\n| J | mechanical equivalent of heat (778 ft-lb/Btu) |\n| K | friction coefficient |\n| M | Mach number |\n| $\\frac{W}{g}$ | mass flow, slugs per second |\n| $P_t$ | absolute total pressure, pounds per square foot |\n| p | absolute static pressure, pounds per square foot |\n| $p_0$ | absolute barometric pressure, pounds per square foot |\n| q | dynamic pressure, pounds per square foot |\n| R | gas constant, foot-pounds per pound per degree Fahrenheit |\n| $T_t$ | total temperature, degrees Fahrenheit absolute |\n| T | static temperature, degrees Fahrenheit absolute |\n| V | velocity, feet per second |\n| $W_a$ | air flow, pounds per second |\n| $W_f$ | fuel flow, pounds per second |\n| $\\gamma$ | ratio of specific heat at constant pressure to specific heat at constant volume (considered as variable herein) |\n| $\\delta_0$ | ratio of absolute barometric pressure to NACA standard atmospheric pressure at sea level, 2116 pounds per square foot absolute ($p_0/2116$) |\n| $\\eta$ | over-all efficiency |\n| $\\eta_b$ | combustion efficiency |\n| $\\eta_{tc}$ | thermodynamic-cycle efficiency |\n| $\\eta_d$ | diffuser efficiency |\n| $\\eta_p$ | propulsive efficiency |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:20.599272+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 17, "total_pages": 22, "image_filename": "19930085922_p17.jpg", "text": "```markdown\n16\nNACA RM No. L9C23\n\n<!-- Image (108, 125, 921, 888) -->\n\nFigure 7.- Effect of Mach number on the wing-tip helix angle obtained for several angles of attack with various aileron deflections; vertical fins off; $\\delta_{a_r} = 0$.\n```", "timestamp": "2026-07-22T06:44:22.540415+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 21, "total_pages": 52, "image_filename": "19930085911_p21.jpg", "text": "20\nCONFIDENTIAL\nNACA RM E9F22\n\nIn order to facilitate the calculations, graphs were made of\nequations (12), (13), (14), and (15). Wherever possible, tables\nfrom reference 6 are used for equations (2), (3), (4), (5), and\n(8).\n\nREFERENCES\n\n1. Kinghorn, George F., and Disher, John H.: Free-Flight Investi-\ngation of 16-Inch-Diameter Supersonic Ram-Jet Unit. NACA\nRM E8A26, 1948.\n\n2. Wilcox, Fred A., and Howard, Ephraim M.: Comparison of Two\nFuels in Bumblebee 18-Inch Ram Jet Incorporating Rake-Type\nFlame Holder. NACA RM E8F11, 1948.\n\n3. Ferri, Antonio, and Nucci, Louis M.: Preliminary Investigation\nof a New Type of Supersonic Inlet. NACA RM L6J31, 1946.\n\n4. Perchonok, Eugene, Sterbentz, William H., and Moore, Stanley H.:\nIndirect Methods for Obtaining Ram-Jet Exhaust-Gas Temperature\nApplied to Fuel-Metering Control. NACA RM E7H27, 1948.\n\n5. Turner, L. Richard, and Lord, Albert M.: Thermodynamic Charts\nfor the Computation of Combustion and Mixture Temperatures at\nConstant Pressure. NACA TN 1086, 1946.\n\n6. The Staff of the Ames 1- by 3-Foot Supersonic Wind Tunnel Sec-\ntion: Notes and Tables for Use in the Analysis of Supersonic\nFlow. NACA TN 1428, 1947.\n\nCONFIDENTIAL\n1152", "timestamp": "2026-07-22T06:44:23.891201+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 46, "total_pages": 60, "image_filename": "19930085862_p46.jpg", "text": "```markdown\n44\nNACA RM No. L9A07\n\n<!-- Image (100, 120, 831, 300) -->\n\nFlap configuration\nOff\nSplit\nDroop and split and fences\nL.e. and split\nL.e. and split and fences\nL.e. and fences\n\n<!-- Image (100, 450, 831, 850) -->\n\nFigure 14.- Rolling- and yawing-moment characteristics for a total\naileron deflection of $30^\\circ$.\n```", "timestamp": "2026-07-22T06:44:25.617969+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 45, "total_pages": 149, "image_filename": "19930083192_p45.jpg", "text": "NACA TN 1976\n41\n\nincreased. Figure 36 shows that for a mass parameter of 10, a 1-percent change in center-of-gravity position represents about a 1-percent change in load factor, whereas tripling the mass parameter to a value of about 30 increases the rate of change in the total wing-load increment 2 percent for each 1-percent shift in center-of-gravity position.\n\nThe tailless airplane for which calculations were made had a mass parameter of about 26, and a $1\\frac{1}{2}$-percent increase in load for a 1-percent rearward movement of the center of gravity for 8 chords was indicated (fig. 38(a)), which is in fair agreement with the curves shown in figure 34 for the conventional airplane. Although the analysis is crude, the results indicate that, for both the conventional and tailless airplanes, the effect of the center-of-gravity position is about the same. Consideration of figure 34 indicates that, although the center of gravity is significant in determining the changes in load on the airplane, the center-of-gravity position is not significant in setting the absolute level of the total load for any particular airplane.\n\nTail volume.- The two middle lines of figure 34 for approximately equal tail volume indicate that this quantity by itself is not of prime importance. Inspection of the results for a gradient distance of 8 chords shows that for equal tail volumes but different tail lengths a variation of 10 percent in the effect of pitch on the applied wing load increment can be obtained for a given static margin. For similar conditions and a gust with a gradient distance of 16 chords, the change in load increment is about 20 percent. A change in tail volume obtained by increasing the tail length and reducing the tail area results in higher loads on the airplane.\n\nThe results shown in figure 34 for the same static margin indicate that the change in tail area does not yield significant variations in load, whereas the change in tail length results in a significant modification of the amount of wing load imposed on the airplane. The data thus indicate that for the maximum alleviation of load by means of pitch of the airplane, a short tail length and large tail area appear to be most beneficial, although the combination will probably lead to an uncomfortable ride due to the violent pitching motion of the airplane. The use of a small area with a short tail length will yield results not too different, although the reduction of damping in pitch may emphasize the pitching oscillation to the detriment of passenger comfort and control of the airplane in continuous rough air.\n\nOn the basis of these analytical results, it is concluded that the tail volume is a secondary factor in establishing the level of wing load. The results indicate that the larger the tail length for a given static margin the greater the load imposed on the wing. For a gradient distance of 8 chords the effect on the total wing load due to the various elements of configuration can vary from 10 to -20 percent.", "timestamp": "2026-07-22T06:44:26.103133+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 25, "total_pages": 29, "image_filename": "19930085899_p25.jpg", "text": "```markdown\n24\n\nM = 1.00\n$\\alpha = 10^\\circ$\nM = 1.10\n$\\alpha = 10^\\circ$\nM = 1.15\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\nWing alone\nWing-fuselage\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 4^\\circ$\n$\\alpha = 4^\\circ$\n\n[Figure: NACA logo]\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n-80 -40 0 40 80\n-80 -40 0 40 80\n-80 -40 0 40 80\n\nTail height, $h_t$, percent semispan\n\nFigure 10.— Concluded.\n\nNACA RM No. 19A21\n```", "timestamp": "2026-07-22T06:44:28.587635+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 12, "total_pages": 47, "image_filename": "19930085919_p12.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 11\n\nof pitching-moment coefficient with lift coefficient up to a lift coefficient of 0.9 than did the drooped-nose flap (figs. 13 and 15). With the split flap deflected $45^\\circ$, the extended-nose flap of full wing span increased the lift coefficient attained before instability more than did the drooped-nose flap of full wing span, but shifted the aerodynamic center forward about 9 percent of the mean aerodynamic chord (figs. 14 and 16). This forward shift of the aerodynamic center due to the extended-nose flap of full wing span was partly alleviated by deflecting the elevon $-20^\\circ$ (the elevon caused a rearward displacement of the aerodynamic center of 3 percent of the mean aerodynamic chord (fig. 16)).\n\nThe drag characteristics of the model with the full-span leading-edge flaps are presented in figure 17. The drooped-nose flap or the extended-nose flap of full wing span reduced the drag coefficients of the model at high lift coefficients, with the split flap either retracted or extended. The extended-nose flap was about twice as effective as the drooped-nose flap in reducing the drag at high lift coefficients.\n\nThe importance of drag at high lift coefficients can be appreciated by considering the sinking speed of an airplane. The variation of lift coefficient with drag coefficient for sinking speeds of 20, 30, and 40 feet per second for an assumed wing loading of 40 pounds per square foot is presented in figure 17. It should be observed that at a lift coefficient of 1.0 the full-span extended-nose flap would decrease the sinking speed of the wing-fuselage combination with the split flap deflected $45^\\circ$ at the wing trailing edge from greater than 40 feet per second to about 30 feet per second. The limiting value of sinking speed recommended in reference 5 is 25 to 30 feet per second.\n\nThe characteristics of the model with the sharp leading edges of 50-percent wing span and of full wing span are shown in figure 18. The sharp leading edges eliminated the increase in longitudinal stability which occurred just above a lift coefficient of 0.3; however, they decreased the lift coefficient attained before the occurrence of longitudinal instability. The addition of the sharp leading edges moved the aerodynamic center forward at low lift coefficients and increased slightly the lift at high angles of attack.\n\nHighest Lift Coefficient Attained\nBefore Longitudinal Instability\n\nAlthough none of the devices eliminated the longitudinal instability, some devices substantially increased the lift coefficient\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:44:34.265369+00:00"}