PDF/ocr_dataset_1050.jsonl

{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 27, "total_pages": 98, "image_filename": "19930086073_p27.jpg", "text": "1.4\n1.2\n1.0\nLift coefficient, $C_L$\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\n0 0 0 0 4 8 12 16 20 24 28 32 36\nAngle of attack, $\\alpha$, deg\n\n$\\triangle$ $\\diamond$ $\\circ$ $\\square$\n44.0 21.2 0 -22.0\nFlap deflection, $\\delta_f$, deg\n\n(a) $C_L$ vs $\\alpha$.\nFigure 4.- Wing alone at 0.0° angle of sideslip with various flap deflections.\n\nNACA RM A9H04\nNACA\n25", "timestamp": "2026-07-22T05:53:56.271333+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 34, "total_pages": 42, "image_filename": "19930085938_p34.jpg", "text": "NACA RM No. L9B04\n\n[Figure: (a) $C_V = 0$; trim = 10.2°.]\n\n[Figure: (b) $C_V = 1.94$; trim = 11.4°.]\n\n[Figure: (c) $C_V = 3.62$; trim = 12.4°.]\n\n[Figure: (d) $C_V = 8.35$; trim = 2.9°; (porpoising).]\n\nFigure 11.— Photographs of single-boom configuration being tested. Full power; gross load coefficient, 3.87.\n\n33", "timestamp": "2026-07-22T05:53:56.673003+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 20, "total_pages": 36, "image_filename": "19930086097_p20.jpg", "text": "18 CONFIDENTIAL NACA RM A9H11\n\nprogressively increased to form wings 3 and 4. The measured values of minimum drag for the revised base shapes at a Mach number of 2.0 are also shown in figure 11. The observed reduction in profile drag as compared to wing 2 clearly indicates the importance of properly designing the airfoil contour near the trailing edge.\n\nMeasurements at Angle of Attack\n\nAirfoil sections composed of circular-arc segments, as illustrated in figure 8 by wings 5, 6, and 7, were used for measuring the characteristics of the blunt-trailing-edge airfoils at angle of attack. The ratio of trailing-edge thickness to maximum thickness for these three wings is 0, 0.5, and 1.0, respectively. The measured lift curves and the drag polars at a Mach number of 1.5 with smooth wing surfaces are shown in figures 12 and 13.¹ The corresponding characteristics at M=2.0 are not shown as they are similar to the results for M=1.5. It may be noted from figure 12 that wing 7, with the fully blunt trailing edge, has approximately a 17-percent greater lift-curve slope than wing 5. The theoretical increase, according to figure 6, is 12 percent. The difference between the theoretical and the measured increase in lift-curve slope is attributed to the difference in viscous effects between blunt- and sharp-trailing-edge airfoils. It is known from the experimental results of Ferri (reference 11) that, even at small angles of attack, the actual lift-curve slope of a sharp-trailing-edge airfoil is less than theory indicates because of flow separation ahead of the trailing edge. At low angles of attack the flow over an airfoil with maximum thickness at the trailing edge would not separate at any point on the airfoil surface. Thus it would be expected that the lift-curve slope of blunt-trailing-edge airfoils would approach the theoretical values more closely than sharp-trailing-edge airfoils. Hence it also would be expected that the measured increase in lift-curve slope due to bluntness would be greater than the theoretical increase calculated from second-order effects in an inviscid flow.\n\nThe effect of bluntness at the trailing edge on the drag polars is illustrated by the curves in figure 13. As in figure 12, the various curves in this figure are for airfoils with a common thickness ratio of 10 percent, and for smooth wing surfaces. The principal experimental\n\n---\n\n¹The lift curves in figure 12 do not pass through the origin of coordinates because the measured data are not corrected for the small stream angle existing in the test section. Although the observed angles for zero lift of the various wings should coincide, these curves show a slight discrepancy because of small constructional differences between the wings.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:53:57.386239+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 40, "total_pages": 42, "image_filename": "19930085914_p40.jpg", "text": "```markdown\nNACA RM A9D25\n\n1.0\n— Cambered and twisted wing ($l_c=.05$), $R=20 \\times 10^6$\n-- Plane wing of ref. 9 ($l_c=.06$), $R=2.35 \\times 10^6$\n\n.8\n\n.6\n\n.4\nM=.20\nM=.60\nM=.80\nM=.89\nM=.93\n\nLift coefficient, $C_L$\n\n.2\nM=.18\nM=.60\nM=.80\nM=.90\nM=.925\n\n0\n\n-.2\n\n-.4\n\n-.6\n0\n.04\n.08\n.12\n.16\n.20\n.24\n.28 for M=.20\nDrag coefficient, $C_D$\n\n(a) $C_L$ vs $C_D$.\n\nFigure 16.— Aerodynamic characteristics at several Mach numbers of the cambered\nand twisted wing and of a wing of identical plan form having no camber or twist.\n\n39\n```", "timestamp": "2026-07-22T05:53:59.034822+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 15, "total_pages": 21, "image_filename": "19930082474_p15.jpg", "text": "NACA TN No. 1799\n13\n\nREFERENCES\n\n1. Stewart, W.: Flight Testing of Helicopters. Jour. R.A.S., vol. 52,\nno. 449, May 1948, pp. 261-292.\n\n2. Sissingh, G.: Contributions to the Problem of Dynamical Stability of\nRotary Wing Aircraft with Articulated Blades (Part III). Reps. and\nTranslations No. 98, British M.A.P. Völkenrode, June 15, 1946.\n\n3. Hohenemser, K.: Longitudinal Stability of the Helicopter in Forward\nFlight. Translation No. F-TS-688-RE, Air Materiel Command,\nAug. 2, 1946.\n\n4. Dingeldein, Richard C., and Schaefer, Raymond F.: Full-Scale\nInvestigation of the Aerodynamic Characteristics of a Typical Single-\nRotor Helicopter in Forward Flight. NACA TN No. 1289, 1947.\n\n5. Gilruth, R. R.: Requirements for Satisfactory Flying Qualities of\nAirplanes. NACA Rep. No. 755, 1943.", "timestamp": "2026-07-22T05:54:04.840325+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 14, "total_pages": 29, "image_filename": "19930093789_p14.jpg", "text": "NACA RM No. E8121 CONFIDENTIAL 13\n\n<!-- Image (258, 136, 860, 806) -->\n\nFigure 2. - Sketch of rotor blade showing turbine dimensions.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:09.186710+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 27, "total_pages": 50, "image_filename": "19930086076_p27.jpg", "text": "NACA RM E9F09\n\n[Figure: Schematic diagram of flame holder 8.]\n\nDimensions:\n- Overall length: 16\"\n- Overall height: 8\"\n- Width of right section: 4\"\n- Vertical spacing between rows: $2\\frac{1}{2}$\"\n- Horizontal spacing between chevrons in row: 1\"\n- Chevron tip-to-tip horizontal distance: $\\frac{1}{2}$\"\n- Chevron height: $\\frac{3}{4}$\"\n- Chevron base width: $\\frac{3}{32}$\"\n- Chevron angle: $\\frac{1}{8}$\"\n- Right section stripe height: $\\frac{1}{16}$\"\n\nFigure 8. - Schematic diagram of flame holder 8.\n\nNACA\n\n25", "timestamp": "2026-07-22T05:54:11.372942+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 41, "total_pages": 51, "image_filename": "19930085962_p41.jpg", "text": "40\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (81, 193, 854, 831) -->\n\nFigure 13.— The variation of lift coefficient with Mach number for various angles of attack of the tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:11.532134+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 60, "total_pages": 96, "image_filename": "19930085880_p60.jpg", "text": "58\nNACA RM No. L9C03\n\nLoad, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\nWetted area, sq ft\n\n(d) $\\tau = 16^\\circ$.\nFigure 18.- Continued.", "timestamp": "2026-07-22T05:54:14.673716+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 16, "total_pages": 34, "image_filename": "19930082472_p16.jpg", "text": "14\nNACA TN No. 1797\n\nTABLE II\nLOCATION OF PRESSURE ORIFICES\n\n| Spanwise Positions¹ of Orifices | |\n| :--- | :--- |\n| **Station No.** | **Percent Semispan** |\n| 1 | 20.9 |\n| 2 | 28.1 |\n| 3 | 41.7 |\n| 4 | 57.4 |\n| 5 | 71.4 |\n| 6 | 85.0 |\n| 7 | 92.5 |\n| 8 | 96.2 |\n\n| Chordwise Positions² of Orifices (on Upper and Lower Surfaces at Each Station³) | |\n| :--- | :--- |\n| **Orifice No.** | **Percent Chord** |\n| 0 | 0 |\n| 1 | .25 |\n| 2 | .50 |\n| 3 | 1.0 |\n| 4 | 1.5 |\n| 5 | 2.5 |\n| 6 | 3.5 |\n| 7 | 5.0 |\n| 8 | 7.5 |\n| 9 | 10.0 |\n| 10 | 20.0 |\n| 11 | 30.0 |\n| 12 | 40.0 |\n| 13 | 50.0 |\n| 14 | 60.0 |\n| 15 | 70.0 |\n| 16 | 80.0 |\n| 17 | 90.0 |\n| 18 | 97.5 |\n\n[NACA logo]\n\n¹Spanwise positions are measured perpendicular to the plane of symmetry.\n\n²Chordwise positions are measured in percent of the windstream chord.\n\n³On station 8, orifices no. 1, 2, 3, 4, 6, 8, 11, 13, 15, 17, and 18 were omitted.", "timestamp": "2026-07-22T05:54:17.458179+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 30, "total_pages": 33, "image_filename": "19930085977_p30.jpg", "text": "```markdown\nM = 0.90\nM = 0.95\nM = 1.00\n\nCONFIDENTIAL\n\n$\\alpha = 10^\\circ$\n$\\alpha = 10^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n.4\n\nWing alone\nWing-fuselage\n\n$\\alpha = 6^\\circ$\n$\\alpha = 6^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 4^\\circ$\n$\\alpha = 4^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\n\n$\\alpha = 0^\\circ$\n$\\alpha = 0^\\circ$\n\n$\\frac{q_{wake}}{q}$\n1.2\n.8\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\nCONFIDENTIAL\n\nFigure 13.— Continued.\n\nNACA RM L9H22\nNACA\n29\n```", "timestamp": "2026-07-22T05:54:18.952426+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 1, "total_pages": 65, "image_filename": "19930082546_p1.jpg", "text": "```markdown\n1871\nC1\nNACA TN No. 1870\n\n# NATIONAL ADVISORY COMMITTEE\n# FOR AERONAUTICS\n\n## TECHNICAL NOTE\nNo. 1870\nC /\n\nENGINEERING DEPT. LIBRARY\nCHANCE-VOUGHT AIRCRAFT\nDALLAS, TEXAS\n\n### FREE-SPACE OSCILLATING PRESSURES NEAR THE TIPS OF\n### ROTATING PROPELLERS\n\nBy Harvey H. Hubbard and Arthur A. Regier\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\n[Figure: NACA logo]\n\nWashington\nApril 1949\n\nAPR 20 1949\n```", "timestamp": "2026-07-22T05:54:19.148448+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 28, "total_pages": 98, "image_filename": "19930086073_p28.jpg", "text": "```markdown\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nLift coefficient, $C_L$\n\n0 0 0 0 .1 .2 .3 .4 .5 .6 .7 .8 .9\nDrag coefficient, $C_D$\n\n$\\triangle$ $\\diamond$ $\\circ$ $\\square$\n44.0 21.2 0 -22.0\nFlap deflection, $\\delta_f$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 4.— Continued.\n\n[Figure: NACA logo]\n\n26\nNACA RM A59E04\n```", "timestamp": "2026-07-22T05:54:21.731171+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 35, "total_pages": 42, "image_filename": "19930085938_p35.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:54:25.180719+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 12, "total_pages": 78, "image_filename": "19930082483_p12.jpg", "text": "10\nNACA TN No. 1807\n\n(1) Pumping loss: Pumping, windage, or fanning loss is caused by the induced circulation of the nonworking gases in the inactive rotor passages. The expression that is used to estimate this loss is similar to the general form listed in reference 4, (equation (7), p. 201) and states that\n\n$$\n\\text{pumping power loss} = \\lambda \\left(\\frac{l}{D}\\right) D^5 \\left(\\frac{N}{1000}\\right)^3 (1-F) \\rho_d \\quad (20)\n$$\n\nwhere\n$\\lambda$ empirical constant, depending upon inactive blade shrouding\n$D$ pitch-line diameter, (ft)\n\nWhen the expression for pumping power loss is corrected for sea-level conditions the equation is\n\n$$\n\\text{pumping power loss} = \\lambda \\left(\\frac{l}{D}\\right) D^5 \\left(\\frac{N}{1000}\\right)^3 \\rho_d (1-F) \\frac{1}{\\delta_1 \\sqrt{\\theta_1}} \\quad (21)\n$$\n\n(2) Driving fluid loss: The power produced by the blading of a full-admission turbine may be expressed as\n\n$$\n\\text{blade power} = W \\frac{\\pi DN}{60} \\left[ (\\bar{c}_{u,1})_{360^\\circ} - (\\bar{c}_{u,2})_{360^\\circ} \\right] \\quad (22)\n$$\n\nThis expression may be written as\n\n$$\n\\text{blade power per blade}\n$$\n\n$$\n= K_{III} \\sum_{r_h}^{r_T} \\bar{\\rho}_1 N (c_{x,1})_{360^\\circ} \\left[ (c_{u,1})_{360^\\circ} - (c_{u,2})_{360^\\circ} \\right] \\quad (23)\n$$\n\nwhere\n\n$$\nK_{III} = \\frac{\\pi^2 D^2 l}{60 b} \\frac{\\bar{\\rho}_1}{\\rho_1}\n$$", "timestamp": "2026-07-22T05:54:25.435235+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 86, "total_pages": 99, "image_filename": "19930082511_p86.jpg", "text": "84\nNACA TN No. 1826\n\n$\\zeta$-plane\n\nA' B' C'\nA B D C\n$i\\eta$ $\\zeta$ $\\xi$ $\\alpha$\n\nz-plane\n$z = e^{\\pi\\zeta}$\n$a = e^{\\pi\\alpha}$\n\nC' B' A' A B D C\n-1 x +1 a\n$iy$\n\n$\\overline{z}_1$\n\n[Figure: NACA logo]\n\nFigure 14.- Physical and transformed spaces for two-dimensional tunnel of unit height with one exit boundary.", "timestamp": "2026-07-22T05:54:27.114723+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 22, "total_pages": 28, "image_filename": "19930086092_p22.jpg", "text": "20\nCONFIDENTIAL\nNACA RM A9F14\n\nPitching-moment coefficient, $C_m$\nDrag coefficient, $C_D$\nLift coefficient, $C_L$\nAngle of sideslip, $\\beta$, deg\n\nAngle of attack, $\\alpha$\n$\\circ$ $0^\\circ$\n$\\square$ $6^\\circ$\n$\\diamond$ $12^\\circ$\n$\\triangle$ $21^\\circ$\n\n(a) $C_m, C_D, C_L$ vs $\\beta$.\nFigure 4. - Aerodynamic characteristics at several angles of attack of the $63^\\circ$ swept-back wing with fuselage. Vertical tail off.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:27.412229+00:00"}
{"citation_id": "19930082474", "source_url": "https://ntrs.nasa.gov/api/citations/19930082474/downloads/19930082474.pdf", "page_number": 16, "total_pages": 21, "image_filename": "19930082474_p16.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:54:32.465998+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 12, "total_pages": 21, "image_filename": "19930091993_p12.jpg", "text": "8\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nThis definition of $\\eta$ is equivalent to the ratio of the effective heating value of the fuel in increasing the enthalpy of the mixture to the actual heating value. According to an unpublished analysis conducted at the NACA Cleveland laboratory, it should be possible to obtain a correlation of the combustion efficiency with other combustion-chamber variables by a relation of the form\n\n$$ \\log (\\eta \\sqrt{T_{T3}}) = a - b/T_{T3} $$\n\nwhere the quantities $a$ and $b$ are only functions of the variable $W_1 \\sqrt{p_{T2}}$. Such a correlation was attempted with the data available by grouping data of nearly equal values of $W_1 \\sqrt{p_{ti}/p_{T2}}$ where $p_{ti}$ is standard atmospheric pressure. This plot shows excellent correlation of the data from $T_{T3}$ of 640° F to the maximum temperature of engine operation (fig. 11). A critical value of the parameter $W_1 \\sqrt{p_{ti}/p_{T2}}$ exists at approximately 3.6 pounds per second, below which there is a sharp drop in combustion efficiency. Because $W_1 \\sqrt{p_{ti}/p_{T2}}$ decreases with increasing altitude, a critical altitude of the engine will occur above which the combustion efficiency of the burner will drop off quite rapidly. During operation of an engine with a burner having such characteristics, no change in the internal aerodynamics of the engine would be noted; the engine would merely require more fuel to operate at the desired speed and temperature.\n\n$F_k(T_{T3})$, which is a function of the turbine temperature, is employed so that $W_3'$ is the weight flow corrected for clearance flow and $W_3 = F_k(T_{T3}) W_3'$ is the actual weight flow through the turbine. The resultant equation in terms of corrected parameters is\n\n$$ \\frac{W_1 n}{\\sigma_{T2} \\delta_{T2}} = \\left[ \\frac{1}{(1+f)} \\frac{\\gamma_2 p_{T3}}{\\gamma_2 p_{T2}} F_k(T_{T3}) \\right] \\frac{W_3' n}{\\sigma_{T3} \\delta_{T3}} \\quad (1) $$\n\nThis equation is more complete than the corresponding one given in reference 4 because of the inclusion of the temperature correction for leakage variation. The compressor-outlet conditions are used to correlate compressor air flow because the compressor-outlet pressure is nearly equal to that at the turbine inlet. The usual equivalent air-flow parameter is multiplied by the equivalent speed to obtain the product $\\sigma_{T2} \\delta_{T2} \\frac{\\gamma_2 p_{T3}}{\\gamma_2 p_{T2}}$ in the gas-state correction factor, thus eliminating the temperature, which would cause a large difference between the compressor and turbine air-flow parameters. These air-flow parameters for the compressor and the turbine are simple combinations of the usual equivalent variables, which are used to correlate compressor and turbine data with only small variations of performance in terms of this parameter over a wide range of gas pressures and temperatures.\n\nIn the engine, the use of variables such as this variable largely eliminates the effects of altitude, ram, or nozzle setting on compressor or turbine performance in terms of equivalent variables. Factors that were neglected are the effect of changes in Reynolds number and heat loss on engine performance. In equation (1), the correction term in brackets is near unity.\n\nThe turbine-power output is absorbed by compressor, bearings, and other accessories. Because enthalpy change is the energy input per pound of gas,\n\n$$ W_1 \\Delta H_c + P_a = W_3 \\Delta H_t $$\n\nwhere $P_a$ is the power absorbed by the bearings and the accessories. In terms of corrected equivalent parameters,\n\n$$ \\frac{W_1 \\Delta H_c}{n \\sigma_{T2} \\delta_{T2}} + \\left[ \\left( \\frac{P_a}{n \\sigma_{T2} \\delta_{T2}} \\right) \\left( \\frac{\\gamma_2 p_{T3}}{\\gamma_2 p_{T2}} \\right) \\right] = \\left[ F_m(T_{T3}) \\frac{\\gamma_2 p_{T3}}{\\gamma_2 p_{T2}} \\right] \\left( \\frac{W_3' \\Delta H_t'}{n \\sigma_{T3} \\delta_{T3}} \\right) \\quad (2) $$\n\nThe torque was used rather than the power in order to avoid directly involving the temperature. The correction terms in brackets introduce a small discrepancy between the compressor and turbine parameters, which have the dimension of torque. The function $F_m(T_{T3})$ is a correction term for torque as affected by clearance expansion at various turbine temperatures. Thus if the prime indicates the value corrected to some standard clearance, the actual turbine-torque parameter is\n\n$$ \\frac{W_3 \\Delta H_t}{n \\sigma_{T3} \\delta_{T3}} = F_m(T_{T3}) \\left( \\frac{W_3' \\Delta H_t'}{n \\sigma_{T3} \\delta_{T3}} \\right) \\quad (3) $$\n\n<!-- Image (90, 426, 480, 610) -->\nFIGURE 11.—Combustion performance of burner.\n\nINTERACTION OF ENGINE COMPONENTS\nSTEADY-STATE JET-ENGINE OPERATION\n\nFor determination of the operation of the compressor-turbine combination in a jet engine, the relations between the compressor and turbine parameters must be used. When no external leakage is assumed, the weight flow through the compressor plus the fuel added is equal to the weight flow through the turbine. Therefore\n\n$$ W_1 = \\frac{1}{(1+f)} W_3 $$\n\nBecause the use of weight flow through the turbine corrected for clearance expansion is desirable, a clearance correction", "timestamp": "2026-07-22T05:54:35.565251+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 15, "total_pages": 29, "image_filename": "19930093789_p15.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:54:36.221648+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 28, "total_pages": 50, "image_filename": "19930086076_p28.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:54:39.782933+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 42, "total_pages": 51, "image_filename": "19930085962_p42.jpg", "text": "NACA RM A9E05 CONFIDENTIAL 41\n\nLift coefficient, $C_L$\n$\\sigma_{approx.}$\n8°\n6°\n4°\n2°\n0°\n-2°\n-4°\n-6°\n-8°\n\nMach number, M\n(b) $\\delta_e$, 4°.\n\nFigure 13. — Continued.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:39.966823+00:00"}
{"citation_id": "19930085914", "source_url": "https://ntrs.nasa.gov/api/citations/19930085914/downloads/19930085914.pdf", "page_number": 41, "total_pages": 42, "image_filename": "19930085914_p41.jpg", "text": "```markdown\n40\n\nLift coefficient, $C_L$\n\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.6\n\n-12 -8 -4 0 4 8 12 16 20 for M=.20\nAngle of attack, $\\alpha$, deg\n\n— Cambered and twisted wing ($l_c$=.05), R=20x$10^6$\n-- Plane wing of ref. 9 ($l_c$=.06), R=2.35x$10^6$\n\nM=.18\nM=.20\nM=.60\nM=.60\nM=.80\nM=.80\nM=.90\nM=.89\nM=.925\nM=.93\n\n[Figure: NACA logo]\n\n(b) $C_L$ vs $\\alpha$.\n\nFigure 16. — Continued.\n\nNACA RM A9D25\n```", "timestamp": "2026-07-22T05:54:40.323045+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 37, "total_pages": 46, "image_filename": "19930085972_p37.jpg", "text": "NACA RM L9B18\n35\n\nDownwash angle, $\\epsilon$, deg\nAngle of attack, $\\alpha$, deg\n\nTail-off lift-curve slope, $(\\frac{dC_L}{d\\alpha})_o$\nHorizontal-tail effectiveness, $\\frac{dC_m}{dC_L}$\n\nNeutral-point and tail-off aerodynamic-center location, $h_p$ and $h_o$, percent $c$ ($C_L=0$)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n(a) Concluded.\nFigure 12.- Continued.", "timestamp": "2026-07-22T05:54:42.291145+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 61, "total_pages": 96, "image_filename": "19930085880_p61.jpg", "text": "NACA RM No. L9C03\n59\n\nLoad, lb\nSpeed\n(fps)\n30\n25\n20\n15\n10\n\nWetted area, sq ft\n\n(e) $\\tau = 20^\\circ$.\n[Figure: NACA logo]\nFigure 18.- Concluded.", "timestamp": "2026-07-22T05:54:44.166416+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 69, "total_pages": 78, "image_filename": "19930082618_p69.jpg", "text": "```markdown\nNACA TN 1945\n67\n\n<!-- Image (278, 104, 786, 904) -->\n\n(b) Airfoils with standard leading-edge roughness.\n\nFigure 16.- Concluded.\n```", "timestamp": "2026-07-22T05:54:44.782290+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 17, "total_pages": 34, "image_filename": "19930082472_p17.jpg", "text": "NACA TN No. 1797\n15\n\n<!-- Image (136, 193, 888, 742) -->\n\nFigure 1.- Geometric characteristics of\n45° swept-forward wing.", "timestamp": "2026-07-22T05:54:47.009581+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 25, "total_pages": 114, "image_filename": "19930086061_p25.jpg", "text": "NACA RM L9J07\n21\n\n9. The lift and pitching moments of the wing with zero-trailing-\nedge sweep agreed remarkably well with those published in NACA\nRM L8G05 for a comparable large-scale wing.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.", "timestamp": "2026-07-22T05:54:48.990515+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 21, "total_pages": 36, "image_filename": "19930086097_p21.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 19\n\nresults for wings 6 and 7, in comparison to wing 5, are summarized in the following table:\n\n| Wing | Bluntness h/t | Change in minimum drag | Increase in lift-curve slope | L/D increase according to equation (23)² | Observed increase in (L/D)max |\n|------|---------------|------------------------|------------------------------|------------------------------------------|-------------------------------|\n| 5    | 0             | 0                      | 0                            | 0                                        | 0                             |\n| 6    | .5            | -8%                    | +13%                         | +7%                                      | +8%                           |\n| 7    | 1.0           | +4%                    | +17%                         | +2%                                      | +3%                           |\n\nAs is evident from these data, the theoretical expectations are again substantiated by the experimental measurements. In particular, the experimental results for wing 7 prove that even in those cases where a blunt-trailing-edge wing may have higher profile drag than a conventional section, it nevertheless is possible for it also to have a higher maximum lift-drag ratio. This, of course, is attributed to the improvement in lift-curve slope, and is evident graphically in figure 13 by the intersection of the two drag polars at a lift coefficient below that which yields maximum lift-drag ratio.\n\nGeneral Discussion\n\nThe foregoing comparison of theory and experiment shows that the theoretical predictions are qualitatively substantiated by the wind-tunnel measurements conducted on airfoils of approximately 10-percent thickness at Mach numbers of 1.5 and 2.0. In accordance with the theoretical calculations it is expected that the improvement in lift of blunt-trailing-edge airfoils over conventional sections will progressively increase as the Mach number is increased beyond about 2.0. Unfortunately, an analogous statement about the reduction in drag cannot be made because of the present limited knowledge about base pressure in two-dimensional flow. As regards thickness-ratio effects, however, simple physical considerations make it apparent that the improvement in lift and drag must approach zero as the airfoil thickness ratio approaches zero.\n\nThe failing of theoretical calculations which indicates that the biconvex and the double-wedge profiles are optimum for specific conditions is, of course, attributed to the assumption of a sharp trailing edge which has been made in previous analyses. The inadequacy of such analyses becomes even more apparent when it is recalled that in the present experiments no attempt has been made either to develop the optimum airfoil shape forward of the base or to use the optimum amount\n\n²The observed change in minimum drag has been used in the evaluation of the increase in L/D from equation (23).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:49.908016+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 29, "total_pages": 98, "image_filename": "19930086073_p29.jpg", "text": "NACA RM A59H04\n\nLift coefficient, $C_L$\n\nPitching-moment coefficient, $C_m$\n\nFlap deflection, $\\delta_f$, deg\n\n-22.0 0 21.2 44.0\n\n(c) $C_L$ vs $C_m$.\n\nFigure 4.—Continued.\n\n27", "timestamp": "2026-07-22T05:54:52.952504+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 36, "total_pages": 42, "image_filename": "19930085938_p36.jpg", "text": "NACA RM NO. L9D04\n\n(a) $C_V = 2.12$; trim = $13.2^\\circ$.\n\n(b) $C_V = 2.92$; trim = $10.4^\\circ$.\n\n(c) $C_V = 6.26$; trim = $13.6^\\circ$.\n\n(d) $C_V = 8.35$; trim = $7.1^\\circ$.\n\nFigure 12.— Photographs of twin-boom configuration being tested. Full power; gross load coefficient, 3.87.\n\n35", "timestamp": "2026-07-22T05:54:53.324754+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 31, "total_pages": 33, "image_filename": "19930085977_p31.jpg", "text": "30\nNACA RM L9H22\n\nCONFIDENTIAL\nM = 1.05\nM = 1.10\n\nWing alone\nWing-fuselage\n\n| | | |\n| :--- | :--- | :--- |\n| | $\\alpha = 10^\\circ$ | |\n| | $\\alpha = 6^\\circ$ | |\n| | $\\alpha = 4^\\circ$ | |\n| | $\\alpha = 0^\\circ$ | |\n\nTail height, $h_t$, percent semispan\nCONFIDENTIAL\nFigure 13.— Concluded.", "timestamp": "2026-07-22T05:54:54.503485+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 87, "total_pages": 99, "image_filename": "19930082511_p87.jpg", "text": "NACA TN No. 1826\n85\n\n$\\zeta$-plane\n\nA'\nB'\nD'\nC'\n\n$i\\eta$\nB\n$\\zeta_1$\nA\nD\nC\n$\\xi$\n$\\infty$\n\n$z_1$\n\n$z$-plane\n$z = e^{\\pi\\zeta}$\n$a = e^{\\pi\\alpha}$\n\nC'\nD'\nB'\nA'\nA\nB\nD\nC\n$-a$\n$-1$\n$\\frac{1}{2}$\n$+1$\n$a$\n$iy$\n\n$\\bar{z}_1$\n\nNACA\n\nFigure 15.- Physical and transformed spaces for symmetrical two-dimensional closed-open-closed tunnel of unit height.", "timestamp": "2026-07-22T05:54:56.422448+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 2, "total_pages": 65, "image_filename": "19930082546_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1870\n\nFREE-SPACE OSCILLATING PRESSURES NEAR THE TIPS OF ROTATING PROPELLERS\n\nBy Harvey H. Hubbard and Arthur A. Regier\n\nSUMMARY\n\nThe theory is given for calculating the free-space oscillating pressures associated with a rotating propeller, at any point in space. Because of its complexity this analysis is convenient only for use in the critical region near the propeller tips where the assumptions used by Gutin to simplify his final equations are not valid. Good agreement was found between analytical and experimental results in the tip Mach number range 0.45 to 1.00 where static tests were conducted.\n\nCharts based on experimental data are included for the fundamental frequencies of two-, three-, four-, five-, six-, and eight-blade propellers and for a range of tip clearances from 0.04 to 0.30 times the propeller diameter. If the power coefficient, tip Mach number, and the tip clearance are known for a given propeller, the designer may determine from these charts the average maximum free-space oscillating pressure in the critical region near the plane of rotation.\n\nAs the tip clearance is decreased, pressures in a region about as wide as one propeller radius are greatly increased. At a constant power the pressure amplitudes of the lower harmonics tend to decrease and the higher harmonics tend to increase with an increase in tip Mach number. The fundamental frequency of pressure produced by a four-blade propeller is essentially independent of tip Mach number in the useful tip Mach number range. At tip Mach numbers near 1.00 so much energy appears in the higher harmonics that the total pressures produced by a two-blade propeller are only slightly greater than those produced by a four-blade propeller at the same tip Mach number and power coefficient.\n\nBlade plan form is shown not to be a significant parameter; however, the nondimensional parameter, tip clearance divided by propeller diameter, is shown to be significant. Pressures in the region ahead of the plane of rotation tend to be out of phase with those behind it. A reflector in the pressure field increases pressures in the plane of its surface by an amount which depends on its shape; a flat surface caused a doubling of the free-space values.", "timestamp": "2026-07-22T05:54:58.277211+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 23, "total_pages": 28, "image_filename": "19930086092_p23.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 21\n\nSide-force coefficient, $C_Y$\n\nYawing-moment coefficient, $C_n$\n\nRolling-moment coefficient, $C_l$\n\nAngle of attack, $\\alpha$\n$\\circ$ $0^\\circ$\n$\\square$ $6^\\circ$\n$\\diamond$ $12^\\circ$\n$\\triangle$ $21^\\circ$\n\nAngle of sideslip, $\\beta$, deg\n\n(b) $C_Y, C_n, C_l$ vs $\\beta$.\n\nFigure 4. — Concluded.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:54:58.490244+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 16, "total_pages": 29, "image_filename": "19930093789_p16.jpg", "text": "NACA RM No. E8I21 CONFIDENTIAL 15\n\n[Figure: Photograph of assembled rotor showing rotating cylindrical shroud formed by blade caps.]\n\nFigure 3. - Photograph of assembled rotor showing rotating cylindrical shroud formed by blade caps.", "timestamp": "2026-07-22T05:55:03.812203+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 29, "total_pages": 50, "image_filename": "19930086076_p29.jpg", "text": "NACA RM E9F09\n27\n\n[Figure: Cutaway view of flame holder 8 after 10 minutes of operation. The image shows two vertical metal components with textured surfaces and rows of small, dark, V-shaped or diamond-shaped indentations along their lengths. A scale bar labeled \"INCHES\" is visible in the lower right corner of the figure area.]\n\nNACA\nC-21869\n7-20-48\n\nFigure 9. - Cutaway view of flame holder 8 after 10 minutes of operation.", "timestamp": "2026-07-22T05:55:06.763069+00:00"}
{"citation_id": "19930085962", "source_url": "https://ntrs.nasa.gov/api/citations/19930085962/downloads/19930085962.pdf", "page_number": 43, "total_pages": 51, "image_filename": "19930085962_p43.jpg", "text": "42\nCONFIDENTIAL\nNACA RM A9E05\n\n<!-- Image (99, 219, 874, 776) -->\n\nFigure 13.— Continued.\n(c) $\\delta_e, 10^\\circ$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:55:09.127331+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 38, "total_pages": 46, "image_filename": "19930085972_p38.jpg", "text": "36\nNACA RM L9B18\n\nNeutral-point and tail-off aerodynamic-center location, $\\eta_p$ and $\\eta_o$, percent $\\bar{c}$ ($\\bar{c}=6$)\nDownwash angle, $\\epsilon$, deg\nTail-off lift-curve slope, $(\\frac{\\partial C_L}{\\partial \\alpha})_o$\nDownwash gradient, $\\frac{\\partial \\epsilon}{\\partial \\alpha}$\nHorizontal-tail effectiveness, $\\frac{\\partial C_m}{\\partial C_L}$\n\n<!-- Image (58, 176, 882, 804) -->\n\n(b) Alternate tail position.\nFigure 12.- Continued.", "timestamp": "2026-07-22T05:55:11.985895+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 13, "total_pages": 78, "image_filename": "19930082483_p13.jpg", "text": "NACA TN No. 1807\n\nK<sub>III</sub> empirical constant for driving-fluid losses to be determined  \nfor particular turbine  \n\nr radius, (ft)  \n\nb total number of rotor blades  \n\nThe subscripts x, h, and T refer to axial, inner-radius (blade root) position, and outer-radius position, respectively.  \n\nAs unbarred symbols, $(c_{x,1})_{360^\\circ}$, $(c_{u,1})_{360^\\circ}$, and $(c_{u,2})_{360^\\circ}$ represent the velocities of the gases at any blade height from root $r_h$ to tip $r_T$. Included in constant $K_{III}$ is $\\frac{\\bar{\\rho}_1}{\\rho_1}$, which was found to be essentially constant over the speed range.  \n\nTherefore, the blade power output of a turbine at full admission may be found by summing up the individual powers developed by all the blades.  \n\n(blade power)<sub>360°</sub>  \n\n$$\n= K_{III} \\sum_{B=0}^{B=b} \\sum_{r_h}^{r_T} \\bar{\\rho}_1 N(c_{x,1})_{360^\\circ} \\left[ (c_{u,1})_{360^\\circ} - (c_{u,2})_{360^\\circ} \\right]\n$$\n\nwhere B is the number of active rotor blades at any instant.  \n\nLosses occur as the result of the scavenging action that must take place as a previously inactive rotor passage enters the active arc of one rotational cycle. The relatively stagnant gases entrapped in the passage during the inactive arc of the cycle are displaced by the recurring active fluid flow. This displacement process requires a rapid acceleration of the displaced gases and a concurrent mixing with the displacing gases that occasion a momentum loss.  \n\nAt the completion of the active arc of the cycle, the reverse occurs and is accompanied by the formation of eddies as the active flow through the channels is reduced and finally cut off entirely.  \n\nThe losses with partial admission resulting from the mixing of active and stagnant gases are manifested as changes in the", "timestamp": "2026-07-22T05:55:12.573519+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 62, "total_pages": 96, "image_filename": "19930085880_p62.jpg", "text": "60\nNACA RM No. L9C03\n\nResistance, lb\nSpeed (fps)\nWetted area, sq ft\n(a) $\\tau = 40^\\circ$.\nFigure 19.- Variation of resistance with wetted area. Model 250B.", "timestamp": "2026-07-22T05:55:13.429004+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 15, "total_pages": 39, "image_filename": "19930093769_p15.jpg", "text": "```markdown\nTABLE II - PERFORMANCE OF 3000-POUND-THRUST TURBOJET ENGINE WITH AN-F-58 FUEL\n\n| Altitude (ft) | Flight Mach number, | Engine speed (rpm) | | Net thrust (lb) | | Specific fuel consumption based on net thrust, (lb/hr)/lb thrust) | | Tail-pipe temperature, (°R) | | Fuel flow (lb/sec) | | NACA fuel number |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | Read | Corrected | Read | Corrected | Read | Corrected | Read | Corrected | Read | Corrected | |\n| 5,150 | 0.2551 | 10,480 | 10,736 | 1343 | 1624 | 1.296 | 1.328 | 1312 | 1377 | 0.484 | 0.599 | 48-210 |\n| 5,200 | .2551 | 11,532 | 11,778 | 1861 | 2245 | 1.177 | 1.202 | 1410 | 1471 | .608 | .750 | 48-210 |\n| 5,200 | .2551 | 12,000 | 12,244 | 2076 | 2505 | 1.141 | 1.164 | 1444 | 1503 | .658 | .810 | 48-210 |\n| 5,200 | .2551 | 12,548 | 12,791 | 2259 | 2725 | 1.185 | 1.208 | 1529 | 1589 | .744 | .915 | 48-210 |\n| 5,100 | .6047 | 10,520 | 10,740 | 1065 | 1285 | 1.477 | 1.508 | 1183 | 1233 | .437 | .538 | 48-210 |\n| 5,100$^a$ | .6047 | 11,540 | 11,758 | 1542 | 1860 | 1.401 | 1.427 | 1301 | 1350 | .600 | .736 | 48-210 |\n| 19,950 | .6072 | 10,536 | 11,366 | 717 | 1578 | 1.370 | 1.477 | 1172 | 1363 | .273 | .648 | 48-206 |\n| 19,900 | .6056 | 11,448 | 12,378 | 1012 | 2210 | 1.316 | 1.423 | 1295 | 1514 | .370 | .874 | 48-206 |\n| 19,860 | .6072 | 11,920 | 12,902 | 1157 | 2546 | 1.356 | 1.468 | 1362 | 1596 | .436 | 1.038 | 48-206 |\n| 19,950 | .6063 | 12,440 | 13,450 | 1308 | 2878 | 1.339 | 1.447 | 1460 | 1707 | .486 | 1.157 | 48-206 |\n| 19,950 | .2462 | 10,500 | 11,264 | 901 | 1976 | 1.264 | 1.356 | 1266 | 1457 | .316 | .744 | 48-206 |\n| 19,950 | .2462 | 11,468 | 12,330 | 1165 | 2555 | 1.185 | 1.275 | 1378 | 1583 | .384 | .905 | 48-206 |\n| 19,900 | .2462 | 11,968 | 12,867 | 1267 | 2759 | 1.199 | 1.289 | 1467 | 1696 | .422 | .988 | 48-206 |\n| 19,950 | .2462 | 12,468 | 13,434 | 1361 | 2985 | 1.214 | 1.308 | 1574 | 1827 | .459 | 1.085 | 48-206 |\n| 19,750 | .8773 | 10,524 | 11,534 | 666 | 1448 | 1.571 | 1.722 | 1074 | 1290 | .291 | .693 | 48-210 |\n| 19,950 | .8748 | 11,460 | 12,620 | 1114 | 2458 | 1.324 | 1.461 | 1177 | 1427 | --- | --- | 48-210 |\n| 19,750 | .8706 | 11,980 | 13,223 | 1349 | 2933 | 1.324 | 1.461 | 1308 | 1594 | .496 | 1.191 | 48-210 |\n| 19,750 | .8690 | 12,468 | 13,843 | 1545 | 3359 | 1.306 | 1.450 | 1391 | 1715 | .560 | 1.353 | 48-210 |\n| 20,500 | .8813 | 10,640 | 11,221 | 742 | 1644 | 1.709 | 1.833 | 1084 | 1247 | .352 | .837 | 48-210 |\n| 19,800 | .8657 | 11,580 | 12,369 | 1125 | 2421 | 1.324 | 1.414 | 1247 | 1423 | .414 | .952 | 48-210 |\n| 20,200$^a$ | .8757 | 12,136 | 12,981 | 1375 | 3025 | 1.255 | 1.342 | 1348 | 1542 | .479 | 1.128 | 48-210 |\n| 19,750 | .8695 | 12,600 | 13,439 | 1565 | 3393 | 1.254 | 1.337 | 1427 | 1623 | .545 | 1.261 | 48-210 |\n| 30,250 | .2442 | 11,616 | 13,037 | 1032 | 3525 | 1.162 | 1.304 | 1413 | 1780 | .333 | 1.276 | 48-206 |\n| 30,100 | .2493 | 12,012 | 13,515 | 1072 | 3620 | 1.146 | 1.289 | 1520 | 1924 | .341 | 1.296 | 48-206 |\n| 30,100 | .2493 | 12,500 | 14,081 | 1151 | 3897 | 1.182 | 1.332 | 1557 | 2103 | .378 | 1.438 | 48-206 |\n| 30,100 | .7050 | 10,468 | 11,894 | 650 | 2195 | 1.237 | 1.406 | 1116 | 1441 | .223 | .857 | 48-206 |\n\nCONFIDENTIAL\n14\nCONFIDENTIAL\nNACA RM No. E8L10a\nNACA\n0107\n```", "timestamp": "2026-07-22T05:55:14.681680+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 18, "total_pages": 34, "image_filename": "19930082472_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:55:16.837896+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 37, "total_pages": 42, "image_filename": "19930085938_p37.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:55:23.745349+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 70, "total_pages": 78, "image_filename": "19930082618_p70.jpg", "text": "68\nNACA TN 1945\n\nSection lift-curve slope per degree, $d\\alpha_l/d\\alpha_o$\n\n.12\n.10\n.08\nO NACA 64-409\n$\\square$ NACA 641-412\n$\\diamond$ NACA 642-415\n$\\triangle$ NACA 643-418\n\n.12\n.10\n.08\nO NACA 641-012\n$\\square$ NACA 641-4212\n$\\diamond$ NACA 641-412\n$\\triangle$ NACA 641-612\n\n.12\n.10\n.08\nO NACA 632-415\n$\\square$ NACA 642-415\n$\\diamond$ NACA 652-415\n$\\triangle$ NACA 662-415\n\n.12\n.10\n.08\nO NACA 0012\n$\\square$ NACA 4412\n$\\diamond$ NACA 4415\n$\\triangle$ NACA 23012\n$\\nabla$ NACA 23015\n\n.5 1.0 2.0 3.0 4.0 5.0 10.0 x $10^6$\nReynolds number, R\n\n(a) Airfoils with smooth surfaces.\n\nFigure 17.- Variation of slope of the section lift curve with Reynolds number for the 15 plain airfoils.", "timestamp": "2026-07-22T05:55:28.021320+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 30, "total_pages": 98, "image_filename": "19930086073_p30.jpg", "text": "1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nLift coefficient, $C_L$\n-.02 0 0 0 0\nRolling-moment coefficient, $C_l$\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\nFigure 4.- Concluded.\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nYawing-moment coefficient, $C_n$\n-.01 0 0 0 0\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nSide-force coefficient, $C_Y$\n0 0 0 0 .04\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\nNACA\n\n28\nNACA RM A9H04", "timestamp": "2026-07-22T05:55:28.562999+00:00"}
{"citation_id": "19930085977", "source_url": "https://ntrs.nasa.gov/api/citations/19930085977/downloads/19930085977.pdf", "page_number": 32, "total_pages": 33, "image_filename": "19930085977_p32.jpg", "text": "```markdown\nNACA RM L9H22\n31\n\nCONFIDENTIAL\n\n[Figure: A schematic diagram of a wing-fuselage configuration with a horizontal tail plane. The diagram shows the relative positions of the wing, fuselage, and tail.]\n\n$\\diamond$ Wing-fuselage $\\alpha = 10^\\circ$\n$\\triangle$ Wing-fuselage $\\alpha = 6^\\circ$\n$\\circ$ Wing alone $\\alpha = 10^\\circ$\n\n$h_t$\n19\n11\n19\n\n[Graph: A plot of $q_{wake}/q$ versus Percent semispan. The y-axis ranges from 0 to 1.2. The x-axis ranges from 0 to 100. The graph contains three data series corresponding to the legend above. A box in the upper left corner of the plot area contains the text \"M = 1.10\". The NACA logo is in the lower right corner of the plot area.]\n\nPercent semispan\nCONFIDENTIAL\n\nFigure 14.- Spanwise dynamic-pressure surveys in region of tail plane for a model with $0^\\circ$ sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil section.\n```", "timestamp": "2026-07-22T05:55:30.226261+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 22, "total_pages": 36, "image_filename": "19930086097_p22.jpg", "text": "20 CONFIDENTIAL NACA RM A9H11\n\nof bluntness at the trailing edge. It is apparent from the present results that extensive experimental work is needed before optimum airfoil shapes can be specified which are satisfactory for engineering purposes. As an example of this, the results show, contrary to the calculations made in reference 12, that at moderate supersonic Mach numbers the optimum airfoil for a given section modulus is not approximately a biconvex section. This is illustrated by the blunt-trailing-edge profile of wing 6 which has about 15 percent greater section modulus than the biconvex profile of wing 5, yet has less drag. (See fig. 13.) As another example, it also may be deduced from the experimental results that the double-wedge section is not close to the optimum for a given airfoil thickness ratio even at moderate supersonic Mach numbers. In particular, a double-wedge profile of the same thickness ratio as wing 4 (9.1 percent) would have about 18 percent less drag than wing 1, since the latter wing has a double-wedge profile of 10-percent thickness ratio. At some Reynolds numbers, however, wing 4 has as much as 30 percent less drag than wing 1 (fig. 10), which would mean about 12 percent less drag than a double-wedge profile of equal thickness ratio.\n\nFrom the viewpoint of immediate practical application, an important engineering problem is that of determining how to avoid large drag increases when considerable bluntness is used on relatively thin airfoils in the low supersonic Mach number range. As was noted earlier, elementary considerations show that it will be difficult to achieve large drag reductions for very thin airfoils since the pressure drag is not a large portion of the profile drag. It is the thin sections, however, which are particularly critical as regards structural difficulties. For example, the depth of the airfoil at the hinge line of a flap is an important structural consideration. In this regard a significant improvement obviously can be obtained even with only a moderately blunt trailing edge. Hence, rather than to concentrate solely on determining optimum airfoil contours for minimum profile drag, it appears to be of equal practical importance to determine how to prevent an appreciable drag increase when employing considerable bluntness on relatively thin airfoil sections.\n\nThe fact that airfoils with blunt trailing edges have higher subsonic drag than conventional sections is a consideration that should be remembered in viewing the possibilities of applying blunt trailing edges to highly swept-back wings. The flow over the outboard regions of a highly swept wing is essentially of the subsonic type, even though the free-stream Mach number is supersonic. In such a case, a blunt-trailing-edge airfoil might increase the profile drag of the outer regions.\n\nIn discussing the possibilities of blunt-trailing-edge airfoils as a practical wing section, incidental advantages can be listed which may be of significance in some designs. For example, the improved structural characteristics near the trailing edge might allow a Fowler-type flap to be used in cases where it could not be used if a conventional airfoil section were employed. The control characteristics at high speeds can also be cited as a possible advantage. Recent experimental investigations\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:55:30.414728+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 88, "total_pages": 99, "image_filename": "19930082511_p88.jpg", "text": "86\nNACA TN No. 1826\n\nz-plane\nC' D' B' A' A B D C\n-a -1 x +1 a\n\nw₁(z)\nD' C' C D\nB' A' A B\n\nw₂(z)\nC' D' D C\nB' A' A B\n\nNACA\n\nFigure 16.- Maps of the functions w₁(z) and w₂(z).", "timestamp": "2026-07-22T05:55:30.733402+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 13, "total_pages": 21, "image_filename": "19930091993_p13.jpg", "text": "ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 9\n\nIn order to evaluate the correction terms, the operating conditions of the engine must be known. Because the effect of these terms is slight, the engine operating conditions need be known only approximately. The engine operation may first be estimated, assuming no corrections, and then the correction terms evaluated. The process may be repeated if desired after the engine operating conditions have been determined with greater accuracy.\n\nBecause the torque parameter $\\frac{W_1 \\Delta H_t}{n \\sigma_{T,2} \\theta_{T,2}}$ is nearly equal to $\\frac{W_3' \\Delta H_t'}{n \\sigma_{T,2} \\theta_{T,2}}$ and the air-flow parameter $\\frac{W_1 n}{\\sigma_{T,2} \\theta_{T,2}}$ is nearly equal to $\\frac{W_3' n}{\\sigma_{T,2} \\theta_{T,2}}$, the compressor and turbine characteristics are plotted in terms of these variables with constant-speed contours. The compressor chart is shown in figure 12. The portions of the curves detached from and to the left of the nearly straight sections represent the region of low air flow and low efficiency, normally in a region of lower air flow than the surge line. A plot of the same variables for the turbine with efficiency contours included is shown in figure 10.\n\nThe approximate simultaneous operation of both the compressor and the turbine is obtained by superimposing the compressor and turbine charts (figs. 10 and 12). Any one point on the compounded chart represents an operating state for both the compressor and the turbine when they are coupled together in the engine. This point has associated with it an equivalent compressor speed $n/\\sqrt{\\theta_{T,1}}$ and an equivalent turbine speed $n/\\sqrt{\\theta_{T,2}}$. Because the compressor and turbine speeds are equal, the ratio of the equivalent speeds is the square root of the corrected temperature ratio:\n\n$$\n\\frac{n/\\sqrt{\\theta_{T,1}}}{n/\\sqrt{\\theta_{T,2}}} = \\sqrt{\\frac{\\theta_{T,2}}{\\theta_{T,1}}}\n\\tag{4}\n$$\n\nIt is thus possible to establish lines of constant equivalent temperature ratio on the compound chart. Any compressor-and turbine-performance variable may then be determined as a function of the compressor equivalent speed and the engine equivalent temperature ratio.\n\nIn order to correct this estimate of engine performance for changes in gas properties, bearing power, leakage losses, and combustion-chamber pressure losses, the altitude of operation is assumed, and the correction terms in equations (1) and (2) are evaluated by means of the approximate correction factors obtained and by making use of the burner characteristics and other auxiliary curves. These corrections applied to the compressor curves give approximate turbine operating conditions required to drive the engine and the resulting curves may be superimposed on the turbine performance charts as were the original uncorrected compressor curves. The process may be repeated if desired for additional accuracy. A set of compressor curves with the corrections introduced and then superimposed on the turbine characteristics is shown in figure 13. The corrections used are based on an inlet-air temperature $T_{T,1}$ of 480° R and an inlet-air pressure $P_{T,1}$ of 1414 pounds per square foot. The heavy line is the surge line. The high-speed range of the compressor is in the region of decreasing turbine efficiency and close to optimum efficiency. A reduction in bearing torque would lower these curves to the center of the high-efficiency region.\n\n[Figure: Torque parameter vs. Air-flow parameter with equivalent compressor speed contours labeled 122, 153, 184, 214, 230, 248, 259, 275, 290 (rps). X-axis: Air-flow parameter, $\\frac{W_1 n}{\\sigma_{T,2} \\theta_{T,2}}$, lb/sec². Y-axis: Torque parameter, $\\frac{W_1 \\Delta H_t}{n \\sigma_{T,2} \\theta_{T,2}}$, ft-pound.]\n\nFIGURE 12.—Compressor performance in parameters for engine computations.\n\n[Figure: Torque parameters vs. Gas-flow parameter with equivalent compressor speed contours (230, 275, 259, 242, 230, 214, 184, 153, 122) and equivalent turbine speed contours (145, 155, 165, 175, 185, 195 rps). Also shows turbine efficiency contours (80, 82, 84, 86, 87, 88, 89, 90, 91, 92, 93, 94, 95, 96, 97, 98, 99) and a Surge line. X-axis: Gas-flow parameter, $\\frac{W_1 n}{\\sigma_{T,2} \\theta_{T,2}}$, lb/sec². Y-axis: Torque parameters, $\\frac{W_1 \\Delta H_t}{n \\sigma_{T,2} \\theta_{T,2}}$, ft-pound.]\n\nFIGURE 13.—Matching chart based on engine experiments. Component performance corrected for losses, clearance expansion, and change in gas properties. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.", "timestamp": "2026-07-22T05:55:32.056112+00:00"}