PDF/ocr_dataset_1052.jsonl

{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 2, "total_pages": 50, "image_filename": "19930082592_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1914\n\nOXIDATION OF TITANIUM CARBIDE BASE CERMALS\nCONTAINING MOLYBDENUM, TUNGSTEN, AND COBALT\n\nBy M. J. Whitman and A. J. Repko\n\nSUMMARY\n\nA number of titanium carbide base cermals were investigated to determine the oxidation-penetration characteristics at various temperatures and exposure periods. Specimens of the various cermals, which were composed of titanium carbide and 5, 10, 20, and 30 percent of molybdenum, tungsten, or cobalt, were oxidized in an air atmosphere at temperatures of $1625^\\circ$, $1785^\\circ$, and $2000^\\circ$ F. A direct-measurement metallographic method was used to determine the depth of oxidation penetration. A metallographic study was made of the oxidation interface and the oxygen-diffusion zone.\n\nWith respect to oxidation penetration, the cermals containing molybdenum were poorest and the tungsten and cobalt cermals were about equal in the time-temperature range for which the data were comparable. The oxides formed on the molybdenum cermals had no protective value in inhibiting further oxidation, whereas the oxides formed on the tungsten and cobalt cermals had protective properties. The cobalt cermals were considered more resistant to over-all oxidation than the tungsten cermals because of their complex, tightly adherent oxide coating.\n\nThe depth of oxide penetration in the temperature range of $1625^\\circ$ to $2000^\\circ$ F with varying compositions up to 30-percent molybdenum, tungsten, or cobalt may be estimated from the oxidation-penetration-time curves.\n\nINTRODUCTION\n\nThe demand for materials that will possess desirable properties at elevated temperatures has resulted in the development of a group of ceramic-metal mixtures known as cermals. These materials combine the excellent refractory characteristics of ceramics with the strength and high-melting-point characteristics of certain metallic elements. The cermals have high strength-to-weight ratios (reference 1) and most of the ceramics considered for use in cermals are easily obtained.", "timestamp": "2026-07-22T05:57:24.325884+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 21, "total_pages": 34, "image_filename": "19930082472_p21.jpg", "text": "NACA TN No. 1797\n19\n\n[Figure: A photograph of a measurement device, a rake, with a ruler marked in inches next to it. The ruler shows markings from 1 to 5 inches. The device has multiple thin probes extending from a base. In the bottom right corner of the image, there is a NACA logo and the identifier A-11428.]\n\nFigure 3.- A typical rake used to survey boundary layers over 45° swept-forward wing.", "timestamp": "2026-07-22T05:57:24.659172+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 15, "total_pages": 21, "image_filename": "19930091993_p15.jpg", "text": "ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 11\n\nFrom this technique for computing engine operation, where the turbine characteristic curves are vertical, indicating turbine choking, the compressor performance in the engine may be accurately predicted independent of errors in the turbine efficiency. The exhaust-pressure ratio is, however, dependent on the turbine efficiency.\n\nOver-all corrected static thrust and specific thrust of the engine were computed from the matching-chart data, with the assumption that there were no losses in the intake or the jet nozzle. The resultant performance curves, including operation at all possible ram pressures and jet-nozzle sizes, are shown in figure 17, which also shows contours of constant compressor speed and constant temperature ratio and the compressor surge line. For operation at a fixed speed, increase of the temperature ratios beyond the limit shown by the surge line causes a sharp drop in thrust and specific thrust. At a temperature ratio of 3, a thrust of 678 pounds can be obtained for a specific thrust of 1.06 pounds thrust per pound fuel per hour and a compressor speed of 290 rps. The maximum operating temperature ratio that can be used at 290 rps without surge is 4, at which condition the engine thrust is 965 pounds and the specific thrust 0.84 pound thrust per pound fuel per hour. The compressor efficiency did not begin to fall off at 290 rps so that engine thrust and specific thrust could probably be improved by operating at speeds higher than 290 rps in the high-temperature and high-thrust range.\n\n[Figure: Sea-level equivalent specific thrust, lb thrust/lb fuel/hr vs Sea-level equivalent thrust, lb. Curves for corrected temperature ratio (γR₁T₁/γR₁T₁) = 2.0, 2.5, 3.0, 3.5, 4.0; equivalent compressor speed n/√θT,1 (rps) = 250, 275, 290, 350; surge line marked.]\n\nFIGURE 17.—Equivalent static performance of turbojet engine. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.\n\nESTIMATED EFFECT OF ENGINE MODIFICATIONS\n\nThe engine data were next applied to the prediction of engine performance with assumed improvements in the components. The burner used in the engine had a peak combustion efficiency of 87 percent. Burners now available show a peak combustion efficiency of 95 percent and a friction coefficient about 40 percent of that of the burner employed in this engine. A burner with these characteristics was assumed to be installed in the engine. Another modification was the assumed installation of a set of low-friction bearings instead of the journal bearings used in the experiments. The turbine matching chart for operation at various engine temperature ratios computed on this basis for an inlet temperature of 480° R and an inlet pressure of 1414 pounds per square foot is shown in figure 18. The essential change is\n\n[Figure: W.M./no.θT,1 (l-pound) vs Gas-flow parameter, W₁n/(σ₁θT,1), lb/sec². Curves for corrected temperature ratio γR₁T₁/γR₁T₁ = 2.0, 2.5, 3.0, 3.5, 4.0, 4.5, 5.0; equivalent turbine speed n/√θT,1 (rps) = 105, 125, 135, 145, 155, 165, 175, 185, 195; torque parameter and turbine efficiency contours labeled.]\n\nFIGURE 18.—Estimated turbine performance at various temperature ratios in engine with assumed modifications. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.\n\nthe reduction in turbine torque required to operate the engine at various speeds. Operating conditions for desirable high speeds are now lowered into the region of peak turbine efficiency. The improvement of the bearings is therefore of special significance in this case because it not only allows more power for the jet but also causes the turbine to operate with higher efficiency. The point at the temperature ratio of 3.5 and a turbine speed n/√θT,1 of 155 rps corresponding to a compressor speed n/√θT,1 of 290 rps is in the region of peak turbine efficiency. The compressor operating conditions are only slightly changed, as shown in figure 19. The\n\n[Figure: Compressor-to-burner-pressure ratio, p₂/p₁ vs Air-flow parameter, W₁n/(σ₁θT,1), lb/sec². Curves for corrected temperature ratio γR₁T₁/γR₁T₁ = 2.0, 2.5, 3.0, 3.5, 4.0; equivalent compressor speed n/√θT,1 (rps) = 250, 275, 290, 345, 385; compressor efficiency contours labeled.]\n\nFIGURE 19.—Compressor performance at various temperature ratios in turbojet engine with assumed modifications. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.", "timestamp": "2026-07-22T05:57:26.064493+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 33, "total_pages": 98, "image_filename": "19930086073_p33.jpg", "text": "```markdown\nNACA RM A9E04\n\n1.4\n1.2\n1.0\nLift coefficient, $C_L$\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\n.12 .08 .04 0 0 0 0 -.04 -.08 -.12 -.16 -.20 -.24 -.28\nPitching-moment coefficient, $C_m$\n\n$\\square$ $\\diamond$ $\\triangle$\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\n\n(c) $C_L$ vs $C_m$.\n\nFigure 5.— Continued.\n\n31\n\n```", "timestamp": "2026-07-22T05:57:28.688500+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 16, "total_pages": 78, "image_filename": "19930082483_p16.jpg", "text": "```markdown\n14\nNACA TN No. 1807\n\n(net power)$_F$ = (ideal power)$_F$ - (rotor-tip leakage loss)$_F$\n- (nozzle and rotor-blading losses)$_F$ - (shaft losses)$_F$\n- (driving-fluid losses)$_F$\n(29)\n\nThis equation corresponds to equation (3) for full-admission turbines.\n\nFrom the preceding discussion,\n\n(net power)$_F$ = F(gross power)$_{360^\\circ}$\n- (shaft losses)$_F$ - (driving-fluid losses)$_F$\n(30)\n\nPower and efficiency estimation. - Equation (30) indicates a procedure for estimating turbine power output at any degree of admission from the turbine output with full admission once the driving-fluid losses have been determined.\n\nOver-all efficiency for a turbine operating with partial admission then, is calculated,\n\n$$ \\eta' = \\frac{(\\text{net power})_F}{(\\text{ideal power})_F} $$\n(31)\n\nEfficiency estimations are based on the assumption that the ratio of the efficiencies of any two degrees of admission is proportional to the ratio of the observed powers per pound of driving fluid developed at these admissions. The equation\n\n$$ \\frac{\\eta'_F}{\\eta'_{360^\\circ}} = \\frac{\\left(\\frac{P_F}{W_F}\\right)}{\\left(\\frac{P_{360^\\circ}}{W_{360^\\circ}}\\right)} $$\n(32)\n\ncompares the efficiency for partial admission with full-admission data.\n\nBecause, as previously discussed, the weight flow is assumed proportional to the degree of admission, equation (32) becomes\n\n1032\n```", "timestamp": "2026-07-22T05:57:29.180888+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 19, "total_pages": 29, "image_filename": "19930093789_p19.jpg", "text": "18\nCONFIDENTIAL\nNACA RM No. EB121\n\n[Figure: Technical drawing of a turbine stator with multiple views and dimensions]\n\nHalf of\ncone angle\n36°\n36°\n14.53\" diam.\n11.62\" diam.\n2 10/4\"\nSection A-A\n\n1.06\"\n3 7/16\"\n20°\nSection B-B x 2\n\nA\nB\nB\nA\n\nNACA\n\nFigure 5. - 70°-cone-angle turbine stator; 26 stator blades.\n\nCONFIDENTIAL\n1031", "timestamp": "2026-07-22T05:57:33.511750+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 66, "total_pages": 96, "image_filename": "19930085880_p66.jpg", "text": "64\nNACA RM No. L9C03\n\n<!-- Image (101, 109, 860, 866) -->\n\n(e) $\\tau = 20^\\circ$.\nFigure 19.- Concluded.\nNACA", "timestamp": "2026-07-22T05:57:36.037070+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 40, "total_pages": 46, "image_filename": "19930085972_p40.jpg", "text": "```markdown\n38\nNACA RM L9B18\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, $\\frac{dC_{m}}{d i_t}$\n\nNeutral-point and tail-off aerodynamic-center location, $h_o$ and $h_{no}$, percent ($\\Delta c_o$)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n- No cutout\n- - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - -", "timestamp": "2026-07-22T05:57:37.686731+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 3, "total_pages": 44, "image_filename": "19930082566_p3.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE NO. 1889\n\nBIAXIAL FATIGUE STRENGTH OF 24S-T ALUMINUM ALLOY¹\n\nBy Joseph Marin and William Shelson\n\nSUMMARY\n\nThe object of this investigation was to determine the fatigue-strength values for 24S-T aluminum alloy when subjected to various ratios of biaxial stresses. The biaxial stresses considered were both tensile. The influence of various ratios of the maximum values of the principal stresses upon the fatigue strength was determined. Fluctuating biaxial tensile stresses were produced by applying a pulsating internal pressure and an axial tensile load to a thin-walled tubular specimen. The maximum and minimum values of the principal stresses were kept in phase. To apply the dynamic loads, a new type of testing machine was designed and constructed.\n\nS-N diagrams for four principal stress ratios were obtained for defining the fatigue strength up to $5 \\times 10^6$ cycles. An attempt was made to compare the test results with a modified maximum-stress theory of failure but poor agreement was found between theory and test results. The test results show that the uniaxial fatigue strength in the transverse direction of the tubular specimens may be about 60 percent of the fatigue strength in the longitudinal direction.\n\nINTRODUCTION\n\nMany machine and structural parts are subjected to stresses that vary in magnitude with time. For example, a connecting rod may be subjected to fluctuating axial stress which varies from a minimum value $\\sigma_{\\min}$ to a maximum value $\\sigma_{\\max}$, as shown in figure 1. The stress variation in figure 1 can be considered to be made up of a completely reversed or variable stress $\\sigma_r$ superimposed upon a steady mean stress $\\sigma_m$. To determine strength of materials under fluctuating or fatigue stresses, a series of tests are made in a fatigue testing machine. In these tests the specimens are subjected to a given mean stress $\\sigma_m$ and to different values of the maximum stress. For each test the number of cycles of stress required to produce rupture of the specimen are determined and a $\\sigma_{\\max}$-N diagram is plotted as shown in figure 2. For a fixed mean stress it is apparent, as shown in figure 2, that the lower the maximum stress\n\n---\n\n¹New temper designation for alloy used: 24S-T4.", "timestamp": "2026-07-22T05:57:40.451602+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 27, "total_pages": 28, "image_filename": "19930086092_p27.jpg", "text": "NACA RM A9F14 CONFIDENTIAL 25\n\nYawing-moment coefficient, $C_n$\n\nAngle of sideslip, $\\beta$, deg\n\n[Figure: Four graphs showing variation of yawing-moment coefficient with angle of sideslip for various rudder deflections. Top-left graph: Angle of attack, 0°; Top-right graph: Angle of attack, 6°; Bottom-left graph: Angle of attack, 12°; Bottom-right graph: Angle of attack, 21°, with legend for Rudder angle, $\\delta_r$: 0°, -10°, -20°, -25°, -30°. All graphs include a dashed line labeled “Tail off”.]\n\nFigure 7. — Variation of yawing-moment coefficient with angle of sideslip for various rudder deflections. 63° swept-back wing-fuselage combination with vertical tail.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:57:41.427040+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 34, "total_pages": 50, "image_filename": "19930086076_p34.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:57:46.029943+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 22, "total_pages": 34, "image_filename": "19930082472_p22.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:57:51.296950+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 25, "total_pages": 36, "image_filename": "19930086097_p25.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 23\n\n13. Turner, Thomas R., Lockwood, Vernard E., and Vogler, Raymond D.:\nAerodynamic Characteristics at Subsonic and Transonic Speeds of a\n$42.7^\\circ$ Sweptback Wing Model Having an Aileron With Finite Trailing-\nEdge Thickness. NACA RM L8K02, 1949.\n\n14. Sandahl, Carl A.: Free-Flight Investigation at Transonic and Super-\nsonic Speeds of the Rolling Effectiveness of Several Aileron\nConfigurations on a Tapered Wing Having $42.7^\\circ$ Sweepback. NACA\nRM L8K23, 1949.\n\n15. Vaccaro, R. J., and Upton J., Jr.: Hinge Moment Tests of V-349 Model\nat a Mach Number of 1.72. Eng. Dept. Rep. No. 5799. Chance Vought\nAircraft, Stratford, Conn., Nov. 21, 1947.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:58:03.359818+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 28, "total_pages": 114, "image_filename": "19930086061_p28.jpg", "text": "```markdown\nTABLE 1.- LOCATION OF WING ORIFICES\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T05:58:05.080914+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 20, "total_pages": 29, "image_filename": "19930093789_p20.jpg", "text": "NACA RM No. EB121 CONFIDENTIAL 19\n\n1.125\"\n20°\n3-H\nSection B-B x 2\nA\n14.63\" diam.\nB\nB\n11.75\" diam.\nSection A-A\nA\nNACA\nFigure 6. - 0°-cone-angle turbine stator; 36 stator blades.\nCONFIDENTIAL", "timestamp": "2026-07-22T05:58:08.809588+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 7, "total_pages": 65, "image_filename": "19930082546_p7.jpg", "text": "6\nNACA TN No. 1870\n\nCertain geometric relations used in the analysis are shown in figure 1. The propeller lies in the zy-plane and the observer is in the xy-plane. The radius of a doublet circle is r. The doublet under consideration is located at the origin of the primed coordinates with angles to observer indicated by $\\delta$, $\\chi$, and $\\nu$. The distance between the observer and the doublet is S. The coordinates of the observer in the primed coordinate system are\n\n$$\n\\begin{aligned}\nx' &= x \\\\\ny' &= y - r \\cos \\theta \\\\\nz' &= -r \\sin \\theta\n\\end{aligned}\n$$\n\nTherefore,\n\n$$\nS = \\sqrt{x^2 + y^2 + r^2 - 2ry \\cos \\theta}\n$$\n\nand\n\n$$\n\\begin{aligned}\n\\cos \\delta &= \\frac{x}{S} \\\\\n\\cos \\chi &= \\frac{y - r \\cos \\theta}{S} \\\\\n\\cos \\nu &= \\frac{-r \\sin \\theta}{S}\n\\end{aligned}\n$$\n\nReference 1 shows that the velocity potential for a given harmonic due to concentrated forces distributed over the propeller disk is given by the following expression:\n\n$$\n\\phi = \\frac{iB}{4\\pi^2\\rho ck} \\int_0^R \\int_0^{2\\pi} \\left[ A(r)e^{i(kct-mB\\theta-\\epsilon_m)}\\cos \\delta + F(r)e^{i(kct-mB\\theta-\\eta_m)}\\cos \\chi \\sin \\theta - \\cos \\nu \\cos \\theta \\right] \\frac{\\partial}{\\partial S} \\left( \\frac{e^{-ikS}}{S} \\right) dr \\, d\\theta\n$$", "timestamp": "2026-07-22T05:58:15.991919+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 67, "total_pages": 96, "image_filename": "19930085880_p67.jpg", "text": "NACA RM No. L9C03\n65\n\n```\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(a) τ = 40.\n```\n[Figure: Graph showing Trimming moment, lb-ft vs Wetted area, sq ft for various speeds (10, 15, 20, 25, 30, 35 fps). Data points are marked with different symbols (circles, squares, triangles, inverted triangles).]\n\nNACA\n\nFigure 20.- Variation of moment with wetted area, Model 250B.", "timestamp": "2026-07-22T05:58:16.102384+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 3, "total_pages": 50, "image_filename": "19930082592_p3.jpg", "text": "```markdown\n2\nNACA TN 1914\n\nCarbide base cermets are receiving particular attention for gas-\nturbine application because they possess strength at elevated temper-\natures and high thermal conductivity. The high thermal conductivity\nis valuable in dissipating heat and is also a factor that contributes\nto good thermal-shock resistance.\n\nIf materials are to be suitable for use as gas-turbine blades,\nthey must also have the capacity to withstand the corrosive effects\nof an air-rich atmosphere at elevated temperature. Quasi-service\nevaluations of turbosupercharger blades of a titanium carbide - cobalt\ncermet indicate that this material has many desirable high-temperature\nproperties but lacks resistance to oxidation (reference 2).\n\nThe research reported herein was conducted at the NACA Lewis\nlaboratory to determine the relative resistance to oxidation in an\nair atmosphere of a group of titanium carbide base cermets at 1625°,\n1785°, and 2000° F and to investigate the types of coating formed.\nTwelve compositions were investigated: 5, 10, 20, and 30 percent of\nthe elements, molybdenum, tungsten, or cobalt, with the balance of\nall cermets consisting of titanium carbide. The cermets are herein-\nafter designated only by the metallic constituent.\n\nThe results presented include the oxidation penetration during\nvarious periods of time, the effect of temperature on oxidation\npenetration, and the effect of change in composition on oxidation\npenetration. The relative merits of molybdenum, tungsten, and cobalt\nas addition elements with regard to oxidation resistance are evaluated\nwithin the temperature range covered. Plots of oxidation-rate con-\nstant against reciprocal of absolute temperature are presented as an\naid in estimating the oxidation penetration of the various materials\nthat will result from exposure to air atmospheres at temperatures\nbetween 1625° and 2000° F.\n\nMetallographic examinations of the oxide coatings were made to\nreveal differences in the oxidation mechanism for the three metallic\naddition elements. Comparisons were made of the tenacity of the\noxides to the cermets.\n\nThe materials were fabricated by Kennametal, Inc. by cold-\npressing fine powder into 1/4-by-1/2-by-4-inch bars that were then\nsintered at suitable temperatures in a controlled atmosphere.\n\nMATERIALS, APPARATUS, AND PROCEDURE\n\nThe cermet specimens used in the oxidation determinations were\ntaken from specimens that had been subjected to a 1600° F modulus-of-\nrupture test in a helium atmosphere. The compositions investigated\n```", "timestamp": "2026-07-22T05:58:19.012854+00:00"}
{"citation_id": "19930086092", "source_url": "https://ntrs.nasa.gov/api/citations/19930086092/downloads/19930086092.pdf", "page_number": 28, "total_pages": 28, "image_filename": "19930086092_p28.jpg", "text": "26\nCONFIDENTIAL\nNACA RM A9F14\n\n<!-- Image (118, 139, 856, 761) -->\n\nFigure 8. - Variation of rudder hinge-moment coefficient with angle of sideslip for various rudder deflections. 63° swept-back wing-fuselage combination with vertical tail.\n\nCONFIDENTIAL\nNACA-Langley - 10-7-49 - 350", "timestamp": "2026-07-22T05:58:23.159172+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 16, "total_pages": 21, "image_filename": "19930091993_p16.jpg", "text": "12\nREPORT 928—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\npoint for a temperature ratio of 3.5 and compressor speed of 290 rps is also in the region of peak compressor efficiency. The confluence of the peak compressor- and turbine-efficiency regions shows that these components are perfectly matched.\n\nThe resultant performance of the modified engine is shown in figure 20. The peak specific thrust in this case is 1.24 pounds thrust per pound fuel per hour. A comparison of figures 20 and 17 shows that for a speed of 290 rps and a temperature ratio of 3.0, the thrust has been increased about 5 percent from 678 to 715 pounds, whereas the specific thrust has been increased 15 percent from 1.06 to 1.22 pounds thrust per pound fuel per hour. At a temperature ratio of 4.0 and a compressor speed $n/\\sqrt{\\theta_{T,1}}$ of 290 rps, the thrust has been increased 9 percent from 965 to 1050 pounds, whereas the specific thrust has been increased 21 percent from 0.84 to 1.02 pounds thrust per pound fuel per hour.\n\ntorque would take place at a higher torque level where the turbine efficiency varied more rapidly than near the region of maximum turbine efficiency.\n\nUnder some conditions, improvement in components might not be reflected in the improvement of over-all engine performance. For example, if an engine with poor bearings is operating at a torque lower than the torque for maximum turbine efficiency, the reduction of torque requirements by improving the bearings would lower the turbine efficiency. Under such circumstances, the over-all performance of the engine might show no change resulting from improvement of the bearings.\n\nACCELERATION OF TURBOJET ENGINE\n\nIn order to observe the interaction of the engine components during acceleration, an estimate of the engine moment of inertia and the torque required for acceleration at the rate of 6.50 rps per second was made. This torque of 38.1 pound-feet was incorporated into the bearing-power correction term $P_b/n\\sigma_T\\theta_{T,1}$ of equation (2). From the form of this equation, the effect of this torque on engine operation is seen to increase inversely with the altitude pressure; acceleration is therefore more difficult at altitude conditions than at sea level. A condition of sea-level pressure was assumed for the computation. The operation of the compressor is shown for the condition of acceleration in figure 21, which shows a slight displacement from the compressor-surge region when compared with the operating condition with no acceleration. For a fixed operating temperature, the acceleration state is equivalent to opening the propulsion nozzle or reducing the back pressure on the turbine in order to obtain the increased turbine power required for acceleration.\n\nHence, if the turbine flow is not choked, the air flow will increase and the back pressure on the compressor will be so lowered that the compressor operating-condition ratio for engine acceleration will be displaced from surge toward the region of lower temperature ratio when compared with\n\n<!-- Image (122, 366, 494, 563) -->\nFIGURE 20.—Static performance of turbojet engine with assumed modifications. Compressor-inlet stagnation temperature and pressure, 480° R and 1414 pounds per square foot, respectively.\n\nThe point at which both the compressor and the turbine are operated at their best efficiency (at a temperature ratio of 3.5 and an equivalent speed of 290 rps) is not the region of best engine efficiency because of decreasing jet efficiency with increasing jet velocities. There will, however, be a flight speed at which the best engine efficiency simultaneously occurs with this optimum matching condition.\n\nThe engine performance estimates for both the unmodified engine and the engine with modifications are somewhat optimistic because the effect of inlet-diffuser and jet-nozzle losses were not taken into account. A reduction of about 4 to 5 percent in thrust and specific thrust would be the right order of magnitude of the correction. Also omitted from the estimate was the power consumption of the oil and fuel pumps that would normally be added to the bearing power consumption when computing engine performance. Although the absolute magnitudes of the performance parameters computed are somewhat optimistic, the effect of engine changes on the improvement of performance is conservative. In the case of the analysis that takes into account the power required for accessories, the reduction in\n\n<!-- Image (527, 648, 945, 888) -->\nFIGURE 21.—Compressor operation with acceleration torque of 38.1 pound-feet at sea-level conditions.", "timestamp": "2026-07-22T05:58:25.088673+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 23, "total_pages": 34, "image_filename": "19930082472_p23.jpg", "text": "NACA TN No. 1797\n\nLeading-edge separation\nTurbulent separation\nTurbulent separation\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\nPitching-moment coefficient, $C_m$\n\nForce data\nPressure data\n\nNACA\n\nFigure 4.—Longitudinal characteristics of 45° swept-forward wing.\n\n21", "timestamp": "2026-07-22T05:58:26.537979+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930082613_p1.jpg", "text": "NACA TN 1938\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1938\n\nMECHANISMS OF FAILURE OF HIGH NICKEL-ALLOY\nTURBOJET COMBUSTION LINERS\n\nBy John W. Weeton\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\n[Figure: NACA logo]\n\nWashington\nOctober 1949", "timestamp": "2026-07-22T05:58:27.161785+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 17, "total_pages": 78, "image_filename": "19930082483_p17.jpg", "text": "NACA TN No. 1807\n\n$\\eta'_{F} = \\frac{1}{F}\\left(\\frac{P_{F}}{P_{s360^{\\circ}}}\\right)\\eta'_{360^{\\circ}}$\n\nPower-Control Parameters\n\nFor a given turbine, power is a function of the independent variables as stated by\n\n$\\text{power} = f\\left(\\operatorname{Re}_{1}, M_{1}, \\gamma, p_{1}', \\frac{p_{1}'}{p_{e}}, N, T_{1}', F\\right)$\n\nwhere\n\nRe Reynolds number\n\nM Mach number\n\nand the subscript e refers to turbine discharge.\n\nOf these parameters, turbine-inlet pressure, turbine pressure ratio, rotor speed, turbine-inlet temperature, and combinations of any two or more of these variables can be used for power control. It is of interest to compare the turbine performance using partial admission as a power control with the turbine performance obtained by using these other parameters to achieve power reduction.\n\nThe individual effects of the normally interacting turbine operating variables may be isolated by allowing these independent parameters to vary one at a time while maintaining the other parameters constant. These methods of power control may then be compared with partial admissions.\n\nEQUIPMENT\n\nTurbine\n\nFor this investigation, the turbine used was obtained from a commercial-type jet-propulsion engine. The turbine components are the turbine-rotor assembly with bearings and bearing supports, the turbine-nozzle assembly and stationary shroud, and the combustion-chamber casing with fuel-nozzle ring.", "timestamp": "2026-07-22T05:58:27.618288+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 26, "total_pages": 36, "image_filename": "19930086097_p26.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:58:27.841150+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 4, "total_pages": 44, "image_filename": "19930082566_p4.jpg", "text": "2\nNACA TN No. 1889\n\nthe greater will be the number of stress cycles that can be applied\nbefore failure occurs. The value of the maximum stress for a given\nnumber of stress cycles, as obtained from the experimental data such\nas figure 2, is called \"the fatigue strength\" of the material. The\nfatigue strength will depend not only upon the number of cycles of\nstress but also upon the value of the mean stress.\n\nMost fatigue tests are made on specimens subjected to simple\nstresses, including fluctuating axial stresses, as described in the\nforegoing paragraph, or fluctuating bending stresses. In machine and\nstructural parts, however, the fluctuating stresses are often not\nsimple or uniaxial stresses, but may be biaxial or triaxial and act\nin more than one direction. There is very little information on the\nfatigue strength of metals subjected to combined stresses. A survey\nof most of the available data is given in reference 1. The purpose\nof this investigation was to obtain the fatigue strength of\n24S-T aluminum alloy when subjected to various ratios of biaxial\nfatigue stresses. Fluctuating biaxial tensile stresses were produced\nby subjecting a tubular specimen to fluctuating axial tension and\nfluctuating internal pressure.\n\nThe project was conducted by the School of Engineering of The\nPennsylvania State College under the sponsorship and with the financial\nassistance of the National Advisory Committee for Aeronautics. The\ntests were conducted in the Combined Stress Laboratory of the Department\nof Engineering Mechanics. Professor K. J. DeJuhasz of the Engineering\nExperiment Station gave valuable suggestions on the design of the\ntesting machine. The testing machine was built by Messrs. M. Aikey,\nH. Johnson, and S. S. Eckley. Messrs. William Stelson and V. L. Dutton,\nresearch assistants, conducted the tests and computed the test data.\nThe administrative direction given by the NACA and the College of\nEngineering and the technical assistance given by the foregoing\nindividuals is greatly appreciated. The testing machine was designed\nby Joseph Marin, who directed the project and prepared this report.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| A | cross-sectional area of tubular specimen, square inches |\n| d | internal diameter of specimen, inches |\n| N | number of stress cycles to failure |\n| p | internal pressure, psi |\n| P | axial tensile load, pounds |\n| p', p'', p''' | maximum, mean, and minimum fluctuating pressure, respectively, psi |", "timestamp": "2026-07-22T05:58:29.661818+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 29, "total_pages": 114, "image_filename": "19930086061_p29.jpg", "text": "NACA RM L9J07\n\n30.0\n\nTunnel\n\nPivot axis\nPivot axis\n\nPitot\ntube\n\nSurvey\nplane\n\n5.3\n4.7\n\n0.5\n\n6.2\n\nCenter of\nentrance cone\n\n30.0\n\n8.0\n0.6\n4.0\n\nSurvey\nplane\n\n14.8\n\nPitot tube\n\n7.6\n\n5.0\n\n(a) Side view.\n\n(b) View looking into entrance cone from rear of model.\n\nNACA\n\nFigure 1.- Sketch of test setup in entrance cone of Langley full-scale tunnel. All dimensions are\nin feet.\n\n25", "timestamp": "2026-07-22T05:58:32.834995+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 21, "total_pages": 29, "image_filename": "19930093789_p21.jpg", "text": "CONFIDENTIAL\n\nLabyrinth-sealing\nair flow\n\nLabyrinth shroud\n0.025-in. radial clearance\nBlade cap\n\nMain air flow\nRotor blade\n\n(a) Labyrinth shroud.\n\n0.025-in. radial clearance\nBlade cap\n\nMain air flow\nRotor blade\n\n(b) Cylindrical shroud.\n\nCylindrical\nshroud\n\nNACA\n\nFigure 7. - Installation of stationary shrouds showing labyrinth-sealing air-flow path and radial clearance.\n\n20\n\nCONFIDENTIAL\n\nNACA RM No. E8121", "timestamp": "2026-07-22T05:58:33.670750+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 73, "total_pages": 78, "image_filename": "19930082618_p73.jpg", "text": "NACA TN 1945\n71\n\n<!-- Image (170, 108, 852, 855) -->\n\n(b) Airfoils with standard leading-edge roughness.\nFigure 18.— Concluded.", "timestamp": "2026-07-22T05:58:34.759535+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 35, "total_pages": 50, "image_filename": "19930086076_p35.jpg", "text": "NACA RM E9F09\n\n$$\n\\begin{array}{c}\n\\text{[Top view: Three zigzag lines with dimensions]} \\\\\n\\begin{array}{c|c|c|c|c}\n & \\frac{1}{4}\" & \\frac{1}{4}\" & \\frac{5}{8}\" & 1\" \\\\\n\\hline\n & & & & \\\\\n\\end{array} \\\\\n\\text{[Side view: Rectangular block with shaded layers]} \\\\\n\\begin{array}{c|c}\n\\frac{1}{16}\" & \\\\\n\\hline\n2\\frac{1}{2}\" & \\\\\n\\hline\n2\\frac{1}{2}\" & \\\\\n\\hline\n & 8\" \\\\\n\\end{array} \\\\\n\\text{[Bottom view: Rectangular block with holes and sloped section]} \\\\\n\\begin{array}{c|c}\n\\frac{1}{2}\" \\text{ holes} & \\\\\n\\hline\n & 20\\frac{1}{4}\" \\\\\n\\hline\n & 4\" \\\\\n\\end{array} \\\\\n\\end{array}\n$$\n\nFigure 12. - Schematic diagram of flame holder 10.\n\n[NACA logo]\n\n33", "timestamp": "2026-07-22T05:58:36.118821+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 41, "total_pages": 46, "image_filename": "19930085972_p41.jpg", "text": "NACA RM L9B18\n39\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, $\\frac{dC_{L_t}}{dC_L}$\n\nNeutral-point and tail-off aerodynamic-center location, $n_p$ and $n_o$, percent c (21-6)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n(e) Sharp-leading-edge wing section\nFigure 12.— Continued.", "timestamp": "2026-07-22T05:58:36.424099+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 34, "total_pages": 98, "image_filename": "19930086073_p34.jpg", "text": "32\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\nLift coefficient, $C_L$\n-.04 -.02 0 0 0 0\nRolling-moment coefficient, $C_l$\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n0 0 0 0 .01\nYawing-moment coefficient, $C_n$\n44.0 21.2 0 -22.0\nFlap deflection, $\\delta_f$, deg\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n-.04 0 0 0 0\nSide-force coefficient, $C_Y$\n-22.0 0 21.2 44.0\nFlap deflection, $\\delta_f$, deg\n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 5- Concluded.\n\nNACA\nNACA RM A59E04", "timestamp": "2026-07-22T05:58:38.833258+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 91, "total_pages": 99, "image_filename": "19930082511_p91.jpg", "text": "NACA TN No. 1826\n\n$\\frac{v}{V} = \\frac{chV}{c_0^2} = \\frac{2ch}{c_0^2}$\n\n[Figure: Graph with two curves labeled .5 and 1.0, plotted against \"Distance from entrance, $\\bar{x}$, units of tunnel height\" on x-axis (0 to 3.2) and $\\frac{v}{V}$ on y-axis (-.2 to .8). Inset diagram shows a channel with height $h$, vortex symbol, and distances labeled $5h$ and $6h$. NACA logo in lower right corner of graph area.]\n\nFigure 19.- Tunnel-induced angle on axis of two-dimensional closed-open-closed tunnel having one exit lip, with vortex at two locations along axis. Length of lower free surface is 1.5 times tunnel height.\n\n89", "timestamp": "2026-07-22T05:58:40.288875+00:00"}
{"citation_id": "19930085938", "source_url": "https://ntrs.nasa.gov/api/citations/19930085938/downloads/19930085938.pdf", "page_number": 41, "total_pages": 42, "image_filename": "19930085938_p41.jpg", "text": "```markdown\n40\n\n14\nTrim, deg\n12\n10\n8\n4\nLoad coefficient, $C_\\Delta$\n3\n2\n1\n0\n0\n1.0\n2.0\n3.0\n4.0\n5.0\n6.0\n7.0\n8.0\nSpeed coefficient, $C_V$\n\n1.4\n7\n1.2\n6\n1.0\n5\nResistance coefficient, $C_R$\nLoad-resistance ratio, $\\Delta_R$\n.8\n4\n.6\n3\n.4\n2\n.2\n1\n0\n0\n\nTail booms\nclear water\nTrim\n$C_R$\n$\\Delta_R$\n$C_\\Delta$\nTransverse spray\nbegins striking\ntail booms\n\nSingle boom ———\nTwin boom - - - - -\n\nNACA\n\nFigure 16.— Minimum stable resistance characteristics for single-boom and twin-boom hulls.\n\nNACA EW NO. 109504\n```", "timestamp": "2026-07-22T05:58:41.266794+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 68, "total_pages": 96, "image_filename": "19930085880_p68.jpg", "text": "66\nNACA RM No. L9C03\n\n[Figure: A graph plotting Trimming moment against Wetted area for various speeds.]\n\nTrimming moment, lb-ft\n18\n16\n14\n12\n10\n8\n6\n4\n2\n0\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(b) $\\tau = 8^\\circ$.\nFigure 20.- Continued.\nSpeed\n(fps)\n30\n25\n20\n15\n10\nNACA", "timestamp": "2026-07-22T05:58:47.312089+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 1, "total_pages": 37, "image_filename": "19930082646_p1.jpg", "text": "FILE COPY\nNO. 6\n\nNACA TN 1980\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nTECHNICAL NOTE 1980\n\nEFFECT OF FOREBODY WARP AND INCREASE IN AFTERBODY LENGTH\nON THE HYDRODYNAMIC QUALITIES OF A FLYING-BOAT\nHULL OF HIGH LENGTH-BEAM RATIO\n\nBy Walter J. Kapryan\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nTHIS DOCUMENT ON LOAN FROM THE FILES OF\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\nLANGLEY AERONAUTICAL LABORATORY\nLANGLEY FIELD, HAMPTON, VIRGINIA\n\nRETURN TO THE ABOVE ADDRESS.\n\nREQUESTS FOR PUBLICATIONS SHOULD BE ADDRESSED\nAS FOLLOWS:\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Washington\n1724 F STREET, N.W.,\nWASHINGTON 25, D.C.\n\n[Figure: NACA logo]\n\nNovember 1949", "timestamp": "2026-07-22T05:58:47.689417+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 24, "total_pages": 34, "image_filename": "19930082472_p24.jpg", "text": "22\nNACA TN No. 1797\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n20.9%\n28.1%\n41.7%\n57.4%\n71.4%\n85.0%\n92.5%\n96.2%\n\nPressure coefficient, P\n-1.2\n-.8\n-.4\n0\n.4\n.8\n\nChordwise station, x/c\n2\n4\n6\n.8\n1.0\n\n(a) $\\alpha=0.1^\\circ$\n\nNACA\n\nFigure 5.—Chordwise pressure distributions\nfor 45° swept-forward wing.", "timestamp": "2026-07-22T05:58:48.321158+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 8, "total_pages": 65, "image_filename": "19930082546_p8.jpg", "text": "NACA TN No. 1870\n\nMaking the substitution for the direction cosines, evaluating $\\frac{\\partial}{\\partial S}\\left(\\frac{e^{-ikS}}{S}\\right)$, and dropping the small phase angles $\\epsilon_m$ and $\\eta_m$ gives\n\n$$\n\\phi = \\frac{-iB}{4\\pi^2\\rho ck} \\int_0^R \\int_0^{2\\pi} \\left[ A(r)e^{i(kct-mB\\theta-kS)}\\frac{x}{S} + F(r)e^{i(kct-mB\\theta-kS)}\\frac{y\\sin\\theta}{S} \\right] \\left( \\frac{ik}{S} + \\frac{1}{S^2} \\right) dr\\ d\\theta\n$$\n\nWhen the concept of an effective radius at which the thrust and torque are assumed to act as in reference 1 is used, and when the following substitutions are also made as in reference 1\n\n$$\nA(r)dr = \\frac{dt}{B}\n$$\n\nand\n\n$$\nF(r)dr = \\frac{dq}{Br}\n$$\n\nthen\n\n$$\n\\phi = \\frac{-ie^{ikct}}{4\\pi^2\\rho ck} \\int_0^{2\\pi} \\left( T_x + \\frac{Q_y\\sin\\theta}{R_e} \\right) \\left( \\frac{ikS_e+1}{S_e^3} \\right) \\left[ \\cos(mB\\theta+kS_e) - i\\sin(mB\\theta+kS_e) \\right] d\\theta\n$$\n\nwhere $R_e$ is an effective radius of the propeller.\n\nThe instantaneous pressure for a given harmonic at any point is given by $p_1 = \\rho \\frac{\\partial\\phi}{\\partial t}$.\n\nHence,\n\n$$\np_1 = \\frac{e^{ikct}}{4\\pi^2} \\int_0^{2\\pi} \\left( T_x + \\frac{Q_y\\sin\\theta}{R_e} \\right) \\left( \\frac{ikS_e+1}{S_e^3} \\right) \\left[ \\cos(mB\\theta+kS_e) - i\\sin(mB\\theta+kS_e) \\right] d\\theta \\tag{1}\n$$", "timestamp": "2026-07-22T05:58:53.500794+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 30, "total_pages": 114, "image_filename": "19930086061_p30.jpg", "text": "26\nNACA RM L9J07\n\n(a) Wing 1.\nA = 3.46\nsta. y, percent b/2\n1 0\n2 16.7\n3 33.3\n4 50.0\n5 66.7\n6 83.3\n7 91.7\n4°\nStation\n2\n3\n4\n5\n6\n7\n60°\n15\n30°\nc = 13.9\n36.0\n\n(b) Wing 2.\nA = 2.31\nSection A-A\nsta. y, percent b/2\n1 0\n2 16.7\n3 33.3\n4 50.0\n5 66.7\n6 83.3\n7 91.7\n4°\nStation\n2\n3\n4\n5\n6\n7\n60°\n15\nc = 20.8\n36.0\n\n(c) Wing 3.\nA = 1.73\nsta. y, percent b/2\n1 0\n2 16.7\n3 33.3\n4 50.0\n5 66.7\n6 83.3\n7 91.7\n4°\nStation\n2\n3\n4\n5\n6\n7\n60°\n15\n30°\nc = 27.7\n36.0\nNACA\n\nFigure 2.- Geometric characteristics of wings tested. All dimensions are in inches.", "timestamp": "2026-07-22T05:58:53.798627+00:00"}
{"citation_id": "19930091993", "source_url": "https://ntrs.nasa.gov/api/citations/19930091993/downloads/19930091993.pdf", "page_number": 17, "total_pages": 21, "image_filename": "19930091993_p17.jpg", "text": "ANALYSIS OF PERFORMANCE OF JET ENGINE FROM CHARACTERISTICS OF COMPONENTS 13\n\nthe steady-state operating condition. For a choked turbine, the compressor operating condition will be the same for acceleration and for steady state with the same compressor speed and engine temperature ratio. In this case, the determination of whether the compressor will operate in the low-efficiency region beyond the surge line when increasing the fuel input for acceleration is a simple matter of examination of the new temperature-ratio point at the compressor speed for the start of acceleration. Less exhaust-nozzle pressure will be available than in the steady-state case for a given speed and acceleration temperature ratio. This fact is shown in figure 22 in which the over-all engine total-pressure ratio for various engine speeds and temperature ratios are shown under the conditions that a torque of 38.1 pound-feet is accelerating the engine. The turbine pressure ratio will increase because of the higher turbine power required for acceleration.\n\nat the compressor inlet. The pressure ratio of the gas generator was first plotted as a function of the mass flow corrected for the gas state at the turbine outlet. The pressure ratio was converted to equivalent isentropic enthalpy drop corrected for gas state at the gas-generator outlet. Contours of constant corrected compressor speed and constant temperature ratio are shown in figure 23. On the basis of such coordinates, turbine performance is independent of speed, as shown in figure 8. A turbine characteristic was computed on this assumption with a critical air flow and pressure ratio, which corresponds in figure 23 to an engine temperature ratio of 4.0 and an equivalent compressor speed of 290 rps. The flow through the first nozzle ring was assumed to be critical at a lower pressure ratio than any other set of blades of the power turbine, with the resultant isentropic enthalpy drop and flow characteristic as shown in figure 23. This curve is merely hypothetical and serves in the absence of a specific power turbine with known characteristics. Normally, each power turbine-speed characteristic would be worked separately and the efficiency would be known at each point of the curve. The intersection of this curve with the gas-generator characteristics gives for each compressor speed a fixed temperature ratio $\\theta_{T,2}/\\theta_{T,1}$, gas flow $W/(\\sigma_{T,1}\\sqrt{\\theta_{T,1}})$, and power output on the assumption of a power turbine efficiency\n\n<!-- Image (38, 352, 447, 625) -->\n\nFIGURE 22.—Over-all pressure ratio of turbojet engine with acceleration torque of 38.1 pound-feet at sea-level conditions.\n\n<!-- Image (508, 435, 883, 853) -->\n\nFIGURE 23.—Superposition of gas-generator and power-turbine characteristics.\n\nOPERATION AS HOT-GAS GENERATOR FOR POWER TURBINE\n\nThe compressor-turbine combination may be regarded as a machine for generating hot gas at a higher pressure than initially available. In the case of the jet engine, this gas is used to propel the airplane by this reaction. Another possible application is the use of this gas for driving a power turbine, which operates independently of the rest of the engine. The operation of such an engine was analyzed by assuming that the exhaust of the turbojet engine was used to drive a power turbine, which expanded the gas to a stagnation pressure equal to that", "timestamp": "2026-07-22T05:58:56.647031+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 18, "total_pages": 39, "image_filename": "19930093769_p18.jpg", "text": "1070\n\nNACA RM No. E8L10a\n\nCONFIDENTIAL\n\n[Figure: Installation of a turbojet engine in a test chamber]\n\nNACA\nC-21669\n6-9-48\n\nFigure 1. - Installation of 3000-pound-thrust turbojet engine with nonburning tail pipe and adjustable-area exhaust nozzle in 10-foot-altitude test chamber.\n\nCONFIDENTIAL\n\n17", "timestamp": "2026-07-22T05:58:59.871727+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 5, "total_pages": 44, "image_filename": "19930082566_p5.jpg", "text": "```markdown\nNACA TN No. 1889\n3\n\nP', P'', P'''\nmaximum, mean, and minimum fluctuating axial\ntension loads, respectively, pounds\n\nR\nprincipal stress ratio\n($\\sigma_2/\\sigma_1 = \\sigma_2'/\\sigma_1' = \\sigma_2''/\\sigma_1''$)\n\n$\\sigma$\nuniaxial tensile stress, psi\n\n$\\sigma_{max}$, $\\sigma_m$, $\\sigma_{min}$, $\\sigma_r$\nmaximum, mean, minimum, and variable uniaxial\ntensile stresses, psi\n\n$\\sigma_1$, $\\sigma_2$\nlongitudinal and transverse biaxial principal\nstresses, psi\n\n$\\sigma_1'$, $\\sigma_1''$, $\\sigma_1'''$\nmaximum, mean, and minimum values of principal\nstress $\\sigma_1$, respectively, psi\n\n$\\sigma_2'$, $\\sigma_2''$, $\\sigma_2'''$\nmaximum, mean, and minimum values of principal\nstress $\\sigma_2$, respectively, psi\n\n$\\sigma_{1t}'$\nfatigue strength for uniaxial longitudinal\ntension, psi\n\n$\\sigma_{1y}$, $\\sigma_{2y}$\nbiaxial yield-strength values, psi\n\n$\\sigma_{1u}$, $\\sigma_{2u}$\nbiaxial nominal ultimate-strength values, psi\n\nDESCRIPTION OF MATERIAL\n\nThe material tested in this investigation was a fully heat-treated\naluminum alloy designated as 24S-T. The material was received in\ntubular extruded form in lengths of 16 feet, with an internal diameter\nof 2 inches and a wall thickness of 1/4 inch. The nominal chemical\ncomposition, in addition to aluminum and normal impurities, consists\nof 4.4 percent copper, 1.5 percent magnesium, and 0.6 percent manganese.\nThe mechanical properties, as furnished by the manufacturer, are:\nTensile strength = 68,000 psi; yield strength (0.2-percent offset) =\n44,000 psi; modulus of elasticity = $10.6 \\times 10^6$ psi; percentage elongation\n(in 2 in.) = 14 percent; and Poisson's ratio = 0.33.\n\nTensile control tests were made on flat specimens machined from the\nwalls of the tubular extrusions. The results of these tests are reported\nin reference 2.\n```", "timestamp": "2026-07-22T05:59:00.048304+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 36, "total_pages": 50, "image_filename": "19930086076_p36.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T05:59:02.998040+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 4, "total_pages": 50, "image_filename": "19930082592_p4.jpg", "text": "NACA TN 1914\n3\n\nconsisted of a titanium carbide base material alloyed with 5, 10,\n20, or 30 percent of cobalt, tungsten, or molybdenum. Chemical\nanalyses of the ceramal compositions investigated showed the presence\nof approximately 6-percent additional tungsten in the form of tung-\nsten carbide in molybdenum and tungsten ceramals (reference 1). In\nthe cobalt ceramals, this contamination is usually less than 1 per-\ncent and the highest percentage analyzed was 2.1 percent of tungsten.\nThe presence of the tungsten carbide was probably due to pickup\nduring milling in ball mills using tungsten carbide balls.\n\nOxidation specimens were prepared by sectioning the modulus-of-\nrupture bars into 1/4-by-1/4-by-1/2-inch rectangular prisms. The\nfaces were so ground that two opposite surfaces (reference faces)\nwere parallel within $\\pm 0.0001$ inch. The distance between reference\nfaces was measured before the oxidation exposure with an accuracy\nof $\\pm 0.0001$ inch.\n\nA resistance furnace with silicon carbide heating elements was\nused to oxidize the specimens in an air atmosphere at temperatures\nof $1625^\\circ$, $1785^\\circ$, and $2000^\\circ$ F. Temperatures controlled by a potenti-\nometer were held to a maximum deviation of $\\pm 15^\\circ$ F. The temperature\nwas checked by means of two chromel-alumel thermocouples placed in\nthe specimen heating zone.\n\nAfter the oxidation period, the oxide coating formed was care-\nfully ground from a surface perpendicular to the reference surfaces.\nDuring grinding, observations were made of the relative resistance\nof the oxide to grinding, the adherence of the oxide to the speci-\nmen, and the presence or absence of striation.\n\nAfter a specimen was ground, the unoxidized cross section of\nthe material was measured by means of a calibrated microscope stage\nand the reticle of a filar eyepiece with an accuracy of $\\pm 0.0004$ inch.\nThe dimensions of the unoxidized area were measured at six points,\nthree evenly spaced along the length and three along the width, and\nthe decrease from the original dimensions averaged for the six meas-\nurements. The microscopic-measurement technique was used because\neven penetration of the oxide formed an easily visible line of\ndemarcation between the unoxidized area and the oxide coating.\nChange in specimen weight during oxidation was considered an inaccurate\nindication of the amount of oxidation because an oxide formed on\nthe molybdenum-ceramal specimens was observed to sublime at elevated\ntemperatures.\n\nRepresentative specimens were mounted in bakelite and polished\nfor metallographic examination of the ceramal, the oxidation inter-\nface, and the oxide over a range of magnification both in the etched\nand unetched condition.", "timestamp": "2026-07-22T05:59:04.500275+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930082613_p2.jpg", "text": "NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTECHNICAL NOTE 1938\n\nMECHANISMS OF FAILURE OF HIGH NICKEL-ALLOY  \nTURBOJET COMBUSTION LINERS\n\nBy John W. Weston\n\nSUMMARY\n\nAn investigation of turbojet combustion-chamber liners from two types of engine was conducted to determine the factors contributing to failure by cracking. Studies were made of \"as-fabricated\", heat-treated, and mechanically finished liners, on liners after service operation, and on liners after accelerated engine runs.\n\nBuckling was produced at or near most cracks by thermal stresses that resulted from over-all temperature gradients and from temperature gradients at individual louvers. Cracks that formed in the buckle were probably caused principally by thermal fatigue of the buckle. Most of the cracks originated at the inner portions of the stress-relieving hole of the louvers in the upper bend of the louver flaps and probably were caused by thermal and mechanical fatigue of the flaps.\n\nCracking may be retarded and liner life prolonged by removing stress raisers produced by punching operations. Cracks are believed to originate both in the grain boundaries and in fissures produced by punching operations. Surface and subsurface scales of appreciable thicknesses were found at edges of metal exposed to the hot gases. Edges of cracks were lined with either or both types of scale and the subsurface scale (internal oxidation) in particular probably contributed to the cracking mechanism by forming stress raisers and subsequently lowering resistance to fatigue.\n\nINTRODUCTION\n\nCombustion-chamber liners used in several types of gas-turbine engine have failed both by cracking and buckling. Cracking is particularly serious because small pieces of metal have been known to break off and be carried into turbine blades causing considerable damage; both cracking and buckling disturb the air flow.", "timestamp": "2026-07-22T05:59:07.430346+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 74, "total_pages": 78, "image_filename": "19930082618_p74.jpg", "text": "72\nNACA TN 1945\n\n3.0\nO NACA 64-409\n□ NACA 641-412\n◇ NACA 642-415\n△ NACA 643-418\n\n2.8\n\n2.6\n\n2.4\n\n2.2\n\n2.0\n\n1.8\n\n1.6\n\n1.4\n\n1.2\n\n1.0\n\n.8\n.5\n\nMaximum section lift coefficients, $c_{l_{max}}$\n\nAirfoils with 20° split\nflaps deflected 60°\n\nPlain airfoils\n\nFlagged symbols and broken lines denote airfoils\nwith standard leading-edge roughness\n\nNACA\n\n1.0 2.0 3.0 4.0 5.0 10.0 x 10⁶\nReynolds number, R\n\nFigure 19.- Variation with Reynolds number of maximum section lift\ncoefficient for four NACA 64-series airfoils of 0.4 design lift\ncoefficient and various thickness.", "timestamp": "2026-07-22T05:59:07.782415+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 25, "total_pages": 34, "image_filename": "19930082472_p25.jpg", "text": "NACA TN No. 1797\n23\n\nUnflagged symbols indicate\nupper surface.\nFlagged symbols indicate\nlower surface.\n\nSpanwise\nstation, 2y/b\n20.9%\n28.1%\n41.7%\n57.4%\n71.4%\n85.0%\n92.5%\n96.2%\n\nPressure coefficient, P\n-1.2\n-.8\n-.4\n0\n.4\n.8\n\nChordwise station, x/c\n2\n4\n6\n8\n10\n\nNACA\n\n(b) $\\alpha=63^\\circ$\n\nFigure 5.—Continued.", "timestamp": "2026-07-22T05:59:08.570677+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 69, "total_pages": 96, "image_filename": "19930085880_p69.jpg", "text": "NACA RM No. L9C03\n67\n\n[Figure: A graph plotting Trimming moment, lb-ft against Wetted area, sq ft. The graph contains multiple curves representing different speeds (10, 15, 20, 25, 30 fps). There is a small inset graph in the upper left corner.]\n\nTrimming moment, lb-ft\nSpeed (fps)\n30\n25\n20\n15\n10\nWetted area, sq ft\n(c) $\\tau = 12^\\circ$.\nFigure 20.- Continued.\nNACA", "timestamp": "2026-07-22T05:59:10.704256+00:00"}
{"citation_id": "19930085972", "source_url": "https://ntrs.nasa.gov/api/citations/19930085972/downloads/19930085972.pdf", "page_number": 42, "total_pages": 46, "image_filename": "19930085972_p42.jpg", "text": "40\nNACA RM L9B18\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, $\\eta_p$ and $\\eta_o$, percent c ($LL_o$)\nLift coefficient, $C_L$\n\nDownwash gradient, $\\frac{d\\epsilon}{d\\alpha}$\nTail-off lift coefficient, $C_{L_o}$\n\n(f) Wing vane.\nFigure 12.- Concluded.", "timestamp": "2026-07-22T05:59:11.597561+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 22, "total_pages": 29, "image_filename": "19930093789_p22.jpg", "text": "NACA RM No. EB121 CONFIDENTIAL 21\n\n70°-cone-\nangle stator\n\nLabyrinth stationary shroud\n\n(a) Configuration 1.\n\nLabyrinth\nstationary\nshroud\n\nCylindrical\nstationary\nshroud\n\n0°-cone-\nangle stator\n\n0°-cone-\nangle stator\n\n(b) Configuration 2.\n\n(c) Configuration 3.\n\n[annotation: no frills !! best]\n\nFigure 8. - Three turbine configurations investigated.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T05:59:13.869847+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 31, "total_pages": 114, "image_filename": "19930086061_p31.jpg", "text": "NACA RM L9J07\n\n[Figure: Close-up view of a wing model mounted on a test apparatus, with a ruler marked in inches visible near the base. A NACA label “L-52216” is affixed to the lower right corner of the image.]\n\n(a) Close-up view of wing 1.\n\nFigure 3.- Wings and test apparatus.\n\n27", "timestamp": "2026-07-22T05:59:14.496217+00:00"}