PDF/ocr_dataset_1055.jsonl

{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 30, "total_pages": 36, "image_filename": "19930086097_p30.jpg", "text": "28\nCONFIDENTIAL\nNACA RM A9H11\n\n[Figure: A photograph showing a model wing mounted inside a large circular wind tunnel test section. A scale is visible on the wall. The NACA logo and identifier A-14229 are in the bottom right corner of the photo.]\n\nFigure 7.—Typical model installation.\n\n[Diagram showing cross-sections of various airfoils labeled 1 through 7:]\n1. Double wedge\n5. Biconvex\n6. Circular arcs (with angle $\\beta$ indicated)\n7. [Triangular wedge shape]\n[Diagram showing a wing planform with span labeled 6c and trailing edge sections labeled 2, 3, 4]\n\n| Wing | Thickness Ratio | h/t | Chord | Trailing-edge Half-angle, $\\beta$ | Relative Section Modulus | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | 10.0% | 0.00 | 2.00in. | 5.7 deg | 1.00 | |\n| 2 | 9.1 | .40 | 2.00 | 3.9 | 1.02 | referred to wing 1 |\n| 3 | 9.1 | .32 | 2.00 | 7.0 | 1.01 | |\n| 4 | 9.1 | .11 | 2.00 | 10.0 | 1.00 | |\n| 5 | 10.0 | .00 | 1.75 | 11.5 | 1.00 | referred to wing 5 |\n| 6 | 10.0 | .50 | 1.75 | 6.9 | 1.15 | |\n| 7 | 10.0 | 1.00 | 1.75 | 0 | 1.00 | |\n\n[Figure 8.—Dimensions and properties of airfoils tested.]\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:03:44.063519+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 74, "total_pages": 96, "image_filename": "19930085880_p74.jpg", "text": "72\nNACA RM No. L9C03\n\n<!-- Image (127, 169, 874, 852) -->\n\nWetted area, sq ft\n(c) $\\tau = 120^\\circ$.\nFigure 21.- Continued.\nNACA", "timestamp": "2026-07-22T06:03:44.549185+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 25, "total_pages": 39, "image_filename": "19930093769_p25.jpg", "text": "24\nCONFIDENTIAL\nNACA RM No. E8L10a\n\nCorrected tail-pipe gas temperature, $C_R$\n\nFuel\n$\\triangle$ AN-F-58\n$\\square$ Gasoline\n\n(a) Altitude, 5000 feet; flight Mach number, 0.\n\n(b) Altitude, 20,000 feet; flight Mach number, 0.60.\n\n(c) Altitude, 20,000 feet; flight Mach number, 0.85.\n\n(d) Altitude, 20,000 feet; flight Mach number, 1.00.\n\nCorrected engine speed, rpm\n\nNACA\n\nFigure 5. - Comparison of corrected tail-pipe gas temperature for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL\n1070", "timestamp": "2026-07-22T06:03:45.045744+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 14, "total_pages": 65, "image_filename": "19930082546_p14.jpg", "text": "NACA TN No. 1870\n\ncondition were obtained from a test airplane. Since the full-scale propeller had 3 blades and operated at much larger power coefficients than the model propellers, no direct comparison could be made. The model data have been extrapolated to the larger power coefficients, however, and interpolations were made at the corresponding tip Mach numbers. The estimates thus obtained are plotted in the following table along with pertinent data from the full-scale tests for comparison:\n\n| $M_t$ | Propeller diameter (ft) | Number of blades | Horse-power | $C_P$ | d/D | $\\bar{p}$, measured (dynes per centimeter²) | $\\bar{p}$, extrapolated from model data (dynes per centimeter²) |\n|-------|--------------------------|------------------|-------------|--------|------|--------------------------------------------|---------------------------------------------------------------|\n| 0.49  | 12.92                    | 3                | 466         | 0.129  | 0.083 | 350                                        | 420                                                           |\n| .49   | 12.92                    | 3                | 466         | .129   | .167 | 240                                        | 280                                                           |\n| .70   | 12.92                    | 3                | 1500        | .135   | .083 | 1500                                       | 1440                                                          |\n| .80   | 12.92                    | 3                | 1500        | .135   | .167 | 1150                                       | 975                                                           |\n\nThus it is seen that model data may be extrapolated to higher values of $C_P$ with a fair amount of accuracy.\n\nHARMONIC ANALYSES OF OSCILLATING PRESSURES\n\nAmplitudes\n\nExperiment.- Data presented thus far have shown the behavior of total oscillating pressures as measured in free space. The subsequent discussion illustrates the behavior of each of the first four harmonics of pressure for a two-blade propeller.\n\nThe effect of power coefficient on the relative amplitudes of the first four harmonics at three different tip Mach numbers in the plane of rotation ($\\frac{x}{D} = 0$) is shown in figure 11. All harmonics are seen to follow the straight-line relationship between $C_P$ and pressure amplitude predicted by equation (2) at $\\frac{x}{D} = 0$. Figure 11(a) shows that for the NACA 4-(5)(08)-03 two-blade propeller the fundamental frequency is predominant at $M_t = 0.75$ and each higher harmonic is smaller in", "timestamp": "2026-07-22T06:03:45.344390+00:00"}
{"citation_id": "19930093789", "source_url": "https://ntrs.nasa.gov/api/citations/19930093789/downloads/19930093789.pdf", "page_number": 29, "total_pages": 29, "image_filename": "19930093789_p29.jpg", "text": "28\nCONFIDENTIAL\nNACA RM No. E8121\n\nLeaning losses not included.\n$$ \\frac{E}{\\Delta_3 h_{t,3}} > \\frac{E}{\\Delta_3 h_{t,3}'} $$\nBrake efficiency, $\\eta = \\frac{E}{\\Delta_3 h_{t,3}'}$\n\nConfiguration\n1\n2\n3\nsee fig 8.\n\nBlade-to-jet speed ratio, $U/V_j$\n\nNACA\n\nFigure 15. - Comparison of configurations 1, 2, and 3 at a total-pressure ratio of 3.00.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:03:46.606655+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 35, "total_pages": 114, "image_filename": "19930086061_p35.jpg", "text": "```markdown\nNACA RM L53J07\n\nVertical distance from tunnel floor, feet\nHorizontal distance from model mount, feet\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:03:48.720118+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 33, "total_pages": 34, "image_filename": "19930082472_p33.jpg", "text": "NACA TN No. 1797\n31\n\n<!-- Image (162, 109, 874, 834) -->\n\nFigure 6. - Boundary-layer growth over six spanwise stations of the 45° swept-forward wing.", "timestamp": "2026-07-22T06:03:53.543217+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 10, "total_pages": 44, "image_filename": "19930082566_p10.jpg", "text": "8\nNACA TN No. 1889\n\nBefore making the internal-pressure adjustments, the oil-pressure line is first fastened to the connection in the lower specimen holder. Various values of maximum- and minimum-pressure readings can be obtained by varying the setting of the eccentric E2 (fig. 8). To apply selected values of the maximum and minimum pressure, a tentative setting of the outer part of the eccentric to the inner part is first made. By means of the hand pump B, the internal-pressure system is filled with oil and the air outlet at the top of the specimen is closed when all the air is expelled from the system. Operation of the hand pump is then continued until the desired minimum internal pressure is reached with the piston of the Bosch pump at the bottom of its stroke. The eccentric drive shaft is now rotated by hand to obtain approximate readings of the minimum and maximum pressures. If the pressure readings are close to the values desired, the motor is switched on so that the actual pressures may be noted. The pressures under dynamic loading with the motor operating are slightly higher than under static loading produced when the shaft is turned by hand. The valve F in the pressure line leading to the pressure gages (fig. 10) is closed during the starting period to avoid shock loading. During a test, however, the valve F is open. Fluctuations of the gages are prevented by a specially designed valve block G, which permits static readings of the maximum pressure on gage H and the minimum pressure on gage L.\n\nThe phase angle between the axial tensile load and the internal pressure can be varied by rotating the inner part of the eccentric E2 relative to the inner part of eccentric E1. To obtain synchronism of the two loads, the eccentrics are adjusted so that the minimum dynamometer axial-load reading is obtained when the piston of the Bosch pump is at the bottom of its stroke.\n\nAfter the desired internal-pressure values are obtained, the axial loading is checked, since the elongation of the specimen produced by the internal pressure reduces slightly the external load produced by the lever. If necessary, the eccentrics E1 and then E2 are again adjusted to give the required values of pressure and axial loads. After the machine has been in operation for about an hour, load adjustments may again be necessary because of changes in temperature of the oil or loosening of the mechanical linkage. To insure correct loading during a test, the loads are checked several times a day. Occasionally it is necessary to add more oil to the system to replace leakage. This is done by means of the hand pump. When the specimen fails, one of the microswitches shuts off the motor, and a record of the number of cycles to failure is recorded by the counter U.\n\nTEST RESULTS\n\nThe fatigue strengths were obtained in this investigation for four principal stress ratios and for ratios of minimum to maximum stress", "timestamp": "2026-07-22T06:03:56.419609+00:00"}
{"citation_id": "19930082646", "source_url": "https://ntrs.nasa.gov/api/citations/19930082646/downloads/19930082646.pdf", "page_number": 6, "total_pages": 37, "image_filename": "19930082646_p6.jpg", "text": "NACA TN 1980\n\nattributed to the increased damping of the oscillations in trim with the extended afterbody. At after positions of the center of gravity, the maximum amplitude of upper-limit porpoising for the modified hull was less than $2^\\circ$. This reduction in the maximum amplitude was also attributed to the effectiveness of the extended afterbody in damping the oscillation in trim.\n\nThe practical center-of-gravity limit for a given elevator deflection is usually defined as that position of the center of gravity at which the amplitude of porpoising becomes $2^\\circ$. Such a plot is presented in figure 8. Since the amplitude of upper-limit porpoising of the modified hull did not reach $2^\\circ$, there is no after limit for this hull. Absence of the after limit is consistent with the results obtained with the extended afterbody (reference 2). This similarity of results would be expected inasmuch as the afterbody has a major influence on upper-limit porpoising which occurs at the after center-of-gravity limit. The forward limit of the modified hull was substantially the same as that of the basic hull, since the shift of the lower trim limit of stability to lower trims was accompanied by a compensating shift of the free-to-trim tracks to lower trims. Because of the elimination of the after limit, the take-off stability characteristics were improved by warping the forebody and extending the afterbody.\n\nLanding Stability\n\nTypical time histories of smooth-water landings made with the modified and basic hulls are presented in figure 9. These and similar time histories were used to determine the maximum amplitudes of oscillation in trim and rise of the center of gravity, which are plotted in figure 10.\n\nThe step was sufficiently deep to prevent skipping of both models at all the contact trims investigated, skipping being defined as the complete emergence of the hull from the water. The maximum trim amplitudes were comparable up to landing trims of $10^\\circ$, except that the cycles of oscillation of the modified hull occurred at lower trims. Above $10^\\circ$, the basic hull encountered greater oscillations. During some landings the modified hull encountered lower-limit porpoising during the run-out after being initially stable, see figure 9(c). This porpoising, which is similar to that noted in reference 2 for the model with the extended afterbody, did not start until the speed decreased to approximately 75 percent of landing speed and would be avoidable through proper control of the elevators. The maximum rise amplitudes were comparable to those of the basic model. The increased amplitudes of trim and rise and the lower-limit porpoising for the model with the extended afterbody (reference 2) apparently were partly compensated by the beneficial effect of warped forebody (reference 1). In general,", "timestamp": "2026-07-22T06:03:57.092310+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 43, "total_pages": 50, "image_filename": "19930086076_p43.jpg", "text": "NACA RM E9F09\n41\n\n[Figure: A photograph of a damaged metal flame holder assembly. The top surface is warped and discolored. The structure has several internal vertical supports. In the background, a metal plate with four circular holes is visible.]\n\nNACA\nC-22396\n10-8-48\n\nFigure 16. - Flame holder 12 after 47 minutes of operation.", "timestamp": "2026-07-22T06:04:03.714411+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 98, "total_pages": 99, "image_filename": "19930082511_p98.jpg", "text": "96\nNACA TN No. 1826\n\n.25\n.20\n.15\n.10\n.05\n$\\delta$ 0\n-.05\n-.10\n-.15\n-.20\n-.25\n\n-.8 -.4 0 .4 .8 1.2 1.6 2.0\n$\\xi$\n\nOpen tunnel\nMethod presented\nMethod of reference 3\nClosed tunnel\nNACA\n\nFigure 26.—Comparison of the results of the present paper with those of reference 3 for $q=-1$, together with those for open and closed circular tunnels.", "timestamp": "2026-07-22T06:04:04.756897+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 39, "total_pages": 98, "image_filename": "19930086073_p39.jpg", "text": "NACA RM A59E04\n\nLift coefficient, $C_L$\n\n|     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     |     | 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{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 2, "total_pages": 26, "image_filename": "19930085485_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:04:06.451830+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930082613_p8.jpg", "text": "NACA TN 1938\n7\n\nstraight paths at this point and turn either away from the center\nline of the louver or parallel to the center line, as shown in fig-\nure 3(a), point 2. The cracks proceed in a slightly crooked path\nbut continue in the same general direction, away from or parallel to\nthe center line, (fig. 3(a), point 3). In extreme cases, cracks\nextended as far as the air-intake holes. Warping or buckling takes\nplace at or near almost every crack. In many cases, the warpage\nforms either a ridge or a groove between the stress-relieving holes\nof the louvers and the nearest air-intake holes (fig. 3(c)). Buck-\nling is most serious in the second row of louvers from the intake\nend (figs. 1 and 4). The louver flaps (fig. 3(a)) do not warp.\n\nSome of the cracks were jagged and widely separated and looked\nlike tears whereas others were minute. The smaller cracks were\nusually much more crooked than the larger ones. Most of the cracks\nwere located in the second and the third row of louvers from the\nintake end (figs. 1 and 4).\n\nSurface examination and chemical analysis. - Inner surfaces of\nthe liners contained carbonaceous deposits, which extended from the\nintake end downstream for approximately 5 inches in type-A liners\nand 8 inches in type-B liners. The thickness of the deposits varied,\nsome large areas having 1/16-inch coatings and a few others having\ngreater thicknesses. A very few localized areas showed green oxide\nspots in the carbonaceous zone. The National Bureau of Standards\nfound the carbonaceous material to be approximately 70-percent carbon\nand 0.3-percent sulfur by microcombustion and combustion analyses,\nrespectively. The spectrographic analysis by the National Bureau\nof Standards did not reveal any significant results. The sulfur\npicked up by the carbonaceous deposits is not believed to contribute\nappreciably to the cracking mechanism because many cracks formed in\nareas away from the carbonaceous deposits.\n\nThermal gradients in liners. - Variation in color of the surface\nscale permitted estimation of the metal temperature. A dull gray-\nblack coating was found on the hottest portions, particularly the\nzone between the first and the third row of louvers. The louvers\nand the adjacent metal 1 or 2 inches upstream were colored dull gray-\nblack whereas immediately downstream were cool areas relatively\nunoxidized and only slightly discolored (fig. 4). Thus, a drastic\ntemperature gradient existed at the individual louvers. Upstream-\ndownstream temperatures at the louvers of type-B liners were measured\nin another NACA investigation and differences of as much as 600° F\nwere found to occur in the second and third rows from the intake end.", "timestamp": "2026-07-22T06:04:08.578248+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 4, "total_pages": 28, "image_filename": "19930082703_p4.jpg", "text": "```markdown\n2\nNACA TN 1983\n\nresults, consisting of time histories of forward-flight maneuvers and\ncorresponding pilot's opinions for each of the three configurations, are\ngiven herein. There is reason to believe that the criterions for heli-\ncopter flying-qualities requirements should remain flexible for some\ntime; however, the indications of the present study together with\nrelated results of reference 1 are interpreted in the form of tentative\n(and incomplete) flying-qualities requirements.\n\nThe stability and control characteristics during recovery from a\ndisturbed flight condition have thus far been found to be more critical,\nin respect to safety, than the immediate effects of a disturbance or\nuncontrolled-for divergence. (See reference 1.) Accordingly, primary\nconsideration is given in this paper to the longitudinal characteristics\nas revealed by pull-ups, which are representative of recovery\ncharacteristics.\n\nResults of theoretical calculations for pull-up maneuvers of two\nhelicopters are also given as a means of indicating whether the charac-\nteristics noted can be theoretically predicted.\n\nMost of the flying represented was done by one NACA test pilot who\nhad had experience in judging the flying qualities of various aircraft.\nEach of the three configurations, however, was also flown to a more\nlimited extent by another NACA pilot, whose impressions on all principal\npoints proved to be the same as those of the first pilot and, hence, are\nnot enumerated separately.\n\nCONFIGURATIONS TESTED\n\nFor convenience, the three configurations are designated heli-\ncopter A, helicopter B, and helicopter C. The object of the investiga-\ntion was to correlate time-history measurements of flight character-\nistics with pilot's reactions to these flight characteristics. Changes\nin details such as center of gravity or the addition or removal of\nexternal equipment are known to affect the comparisons obtained;\nhowever, no attempt is made to discuss the effect of such factors on\nthe comparisons and many details concerning the exact configurations\ntested are omitted.\n\nHelicopter A is a four-place aircraft of about 5,000 pounds gross\nweight and has a single lifting rotor 48 feet in diameter. The general\narrangement is apparent from figure 1.\n\nHelicopter B is the same helicopter with a small fixed horizontal\ntail surface added, this change being sufficient to provide a new\n```", "timestamp": "2026-07-22T06:04:08.918586+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 24, "total_pages": 78, "image_filename": "19930082483_p24.jpg", "text": "22\nNACA TN No. 1807\n\nFor this turbine,\n$$C = 0.0046 \\text{ foot}$$\n$$l = 0.177 \\text{ foot}$$\n$$\\beta_2 = 40.1^\\circ$$\n\nThe leakage weight flow is then 3.86 percent ($K_I$) of the total weight flow.\n\nSubstituting this value of $K_I$ into equation (13) and correcting to sea-level conditions gives\n$$\\text{leakage loss} = \\frac{0.0386 W_a g h'}{\\delta_1 \\sqrt{\\theta_1}} \\quad (38)$$\n\nNozzle and rotor-blading losses. - Nozzle and rotor-blading aerodynamic losses are considered as the differential between ideal and gross power and are not directly determined herein.\n\nDisk-windage loss. - The disk-windage loss is calculated according to equation (15), where\n$$K_{II} = 1.272$$\ndisk power loss\n$$= 1.272 \\left( \\frac{\\rho_d d_h v_h}{\\mu_d} \\right)^{-0.12} \\left( \\frac{N}{1000} \\right)^3 d_h^5 \\rho_d \\frac{1}{\\delta_1^{0.88} \\theta_1^{0.632}} \\quad (39)$$\n\nThe value of 1.272 for the empirical constant is dependent upon the type of disk shrouding and was obtained from the smaller turbine used. This value is considered to be applicable generally to configurations similar to that shown in figure 7. The density term used in the equation represents the average of the densities on both sides of the disk.\n\nBearing loss. - Bearing power loss is calculated from equation (16). The weight flow of oil was determined from the", "timestamp": "2026-07-22T06:04:12.222320+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 75, "total_pages": 96, "image_filename": "19930085880_p75.jpg", "text": "NACA RM No. L9C03\n73\n\nDraft, ft\n.64\n.56\n.48\n.40\n.32\n.24\n.16\n.08\n0\n\nSpeed\n(fps)\n10 O\n15 □\n20 ◇\n25 △\n30 ▽\n\n.05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n(d) $\\tau = 16^\\circ$.\nFigure 21.- Continued.\nNACA", "timestamp": "2026-07-22T06:04:12.683743+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 3, "total_pages": 62, "image_filename": "19930082918_p3.jpg", "text": "```markdown\n2\nNACA TN 1940\n\nThese findings, however, should not be considered generally\napplicable to long-time strength, to strength at other temperatures,\nor to the same alloy in other conditions of treatment. In order to\ndevelop a general theory of high-temperature strength, additional\ndata on other alloys and on the same alloy in the other conditions\nwill have to be gathered.\n\nINTRODUCTION\n\nThis report covers the first part of a continuing investigation\ninto the fundamental behavior at high temperatures of austenitic\nalloys designed for use in aircraft propulsion systems. Previous\ninvestigation (see reference 1, for example) has shown in part the\neffects of prior processing and heat-treatment conditions on the\nhigh-temperature properties of low-carbon N-155 alloy and various\nother austenitic high-temperature alloys. In order to develop\nbetter and practicable alloys on a scientific basis, to utilize\ncritical materials to the fullest possible extent, and to point\nout logical methods of production control for uniform properties,\nmore knowledge must be gained of the fundamental reasons for the\nhigh-temperature behavior of austenitic high-temperature alloys.\n\nThe basic assumption of the investigation was that the behavior\nof certain alloys at high temperature is dependent on their micro-\nstructure and the lattice conditions of the matrix. The experimental\nprogram was therefore designed to measure first these two charac-\nteristics of one alloy, low-carbon N-155, as influenced by heat\ntreatment, chemical composition, and exposure to temperature and\nstress. Optical- and electron-microscope techniques and separation\nand analysis of microconstituents were used to define structural\nconditions. The lattice conditions of the matrix, particularly the\nstrains present, were measured by X-ray diffraction techniques.\nSecond, the creep and rupture properties, corresponding to these\nstructural conditions, were established. Third, these data were\nthen correlated and interpreted using the fundamentals of physics\nof solids and plastic flow to as great an extent as possible.\n\nIt was felt that by concentrating at first wholly on one\nparticular alloy the thorough understanding thus gained could be\nbest generalized for extension to other alloys. The experimental\nvariations which necessarily had to be covered require a prohibitive\namount of work for even two alloys. Low-carbon N-155 alloy was\nchosen primarily because it was a representative alloy of a type\nimportant in the aircraft propulsion field. In addition there was\na considerable background of experimental data for this alloy which\nwould be of value to the investigation.\n```", "timestamp": "2026-07-22T06:04:14.037704+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 31, "total_pages": 36, "image_filename": "19930086097_p31.jpg", "text": "NACA RM A9H11 CONFIDENTIAL 29\n\n.048\nWing 1; theoretical wave drag\n+ laminar friction\n.040\n.032\n.024\n.016\n.008\nLaminar, $M_\\infty=1.5$\n0 .4 .8 1.2\nR, millions\nFigure 9.—Profile drag of wings 1 and 2; $M_\\infty=1.5$, smooth surfaces.\n\n.032\nWing 1; theoretical wave drag\n+ laminar friction\n.024\n.016\n.008\nLaminar, $M_\\infty=2.0$\n0 .4 .8 1.2\nR, millions\nFigure 10.—Profile drag of wings 1, 2, 3, and 4; $M_\\infty=2.0$, smooth surfaces.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:04:15.722063+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 10, "total_pages": 50, "image_filename": "19930082592_p10.jpg", "text": "NACA TN 1914\n\nthat the line is invisible in the oxide coating of the 20- and  \n30-percent-molybdenum ceramal oxidized at $1625^\\circ$ F because the increased  \nquantity of MoO₃ results in the extension of the MoO₃ layer to  \nthe outer surface of the oxide layer, the low oxidation rate in these  \ntwo cases may be explained. The thicker MoO₃ layer would result in  \nslower diffusion of oxygen inward and slower oxidation. At high  \ntemperatures, faster sublimation of the MoO₃ would minimize  \nthis protective action.\n\nThe type of oxidation penetration encountered in the tungsten  \nceramals is illustrated in figure 12. The penetration is regular  \nand the line of demarcation is even. The oxide coating had a  \nyellow-orange color and was easily distinguished from the gray  \nmetallic ceramal. Oxide coatings formed on 5- and 10-percent-  \ntungsten ceramals in a series of striations (fig. 13). When the  \nceramals contained 20 or 30 percent of tungsten, the oxide formed was  \na homogeneous layer with no striations (fig. 12). The appearances of  \nstriated coatings on the ceramals of low tungsten content can be  \nexplained by consideration of the specific molecular volume of the  \noxide formed and the tenacity of the oxide. With low tungsten con-  \ntent, the expanding force of the WO₃ formed is not great enough  \nto break the oxide from the body until a critical thickness of oxide  \nforms. With 20-percent or more tungsten, the volume of WO₃ formed  \nis sufficient to break the TiO₂ and WO₃ away as fast as they  \nform. In figure 5, the variation in oxidation-rate constant with  \ntemperature does not conform with the normal exponential curve. At  \n$2000^\\circ$ F, the oxidation-rate constants of the tungsten ceramals are  \nof approximately the same magnitude. The oxidation-rate constants  \nfor the 20- and 30-percent tungsten ceramals at $1785^\\circ$ F are less  \nthan at $1625^\\circ$ F, whereas at these temperatures the oxidation-rate  \nconstants of the 5- and 10-percent-tungsten ceramals remain prac-  \ntically unchanged.\n\nSeveral possible factors are available to help explain the  \nanomalous oxidation behavior of the tungsten ceramals. The oxide  \nWO₃ undergoes transformation between $1560^\\circ$ and $1650^\\circ$ F (refer-  \nence 8). This transformation results in a noticeable decrease in  \noxidation rate at temperatures above the transformation range  \nbecause of a decrease in solubility of oxygen in the new phase  \n(reference 8). The lowest experimental temperature considered  \nherein is $1625^\\circ$ F, which is within the temperature range suggested.  \nA change in oxygen solubility in WO₃ would result in a change in  \noxidation-rate constant not in keeping with the exponential formula.  \nA further explanation of the wide deviation at $1625^\\circ$ F may be a  \nchange in the actual oxidation-penetration mechanism into the body  \nas the percentage of tungsten is increased. Figures 14 and 15", "timestamp": "2026-07-22T06:04:21.289427+00:00"}
{"citation_id": "19930086076", "source_url": "https://ntrs.nasa.gov/api/citations/19930086076/downloads/19930086076.pdf", "page_number": 44, "total_pages": 50, "image_filename": "19930086076_p44.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:04:26.738010+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 1, "total_pages": 72, "image_filename": "19930085491_p1.jpg", "text": "```markdown\nCopy No. 254\nRM No. A8J04\n\nCONFIDENTIAL\n\nNACA RM No. A8J04\n\n[Figure: NACA logo with wings]\n\n# NACA\n\n## RESEARCH MEMORANDUM\n\nAERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION\n\nEMPLOYING A WING SWEPT BACK $63^\\circ$.- CHARACTERISTICS\n\nAT A MACH NUMBER OF 1.53 INCLUDING EFFECT OF\n\nSMALL VARIATIONS OF SWEEP\n\nBy Robert T. Madden\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officials and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of known loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nJanuary 26, 1949\n\nCONFIDENTIAL\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 113\nEFFECTIVE DATE: MARCH 19, 1967\nWHL\n```", "timestamp": "2026-07-22T06:04:26.938231+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 36, "total_pages": 114, "image_filename": "19930086061_p36.jpg", "text": "```markdown\n32\n\nVertical distance from tunnel floor, feet\n10\n2.5° 2.0° $\\Psi = -35^\\circ$ $\\Psi = 0^\\circ$ $\\Psi = 35^\\circ$ 1.0°\n$\\alpha = 40^\\circ$ $\\alpha = 40^\\circ$ $\\alpha = 40^\\circ$\n2.0°\n9\n8\n1.0° 1.0° Positive angle\n7\n2.5° 2.0° 1.5° 1.0°\n$\\Psi = -35^\\circ$ $\\Psi = 0^\\circ$ $\\Psi = 35^\\circ$\n$\\alpha = 0^\\circ$ 3.0° $\\alpha = 0^\\circ$ $\\alpha = 0^\\circ$\n6\n3.5° Tunnel\n5\n3.5° NACA\n-4 -3 -2 -1 0 1 2 3 4\nLeft Right\nHorizontal distance from model mount, feet\n(a) Pitch angularity.\n\nFigure 5.- Angularity of the air stream in a vertical plane located approximately 7 inches behind\nthe apexes of the wings when $\\Psi = 0^\\circ$ and $\\alpha = 0^\\circ$ and some representative positions of wing 2\nin the surveyed plane.\n\nNACA RM L59D7\n```", "timestamp": "2026-07-22T06:04:27.138680+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 15, "total_pages": 65, "image_filename": "19930082546_p15.jpg", "text": "14\nNACA TN No. 1870\n\namplitude. This order is completely reversed at $M_t = 1.00$ as indicated in figure 11(c). At this speed the fundamental has the smallest amplitude, and the higher-order harmonics are progressively larger. At a tip Mach number of 0.90, which is shown in figure 11(b), the amplitudes are more nearly equal which fact indicates that at this particular speed there is a transition between the two extremes shown in figures 11(a) and 11(c).\n\nThe \"cross over\" phenomenon shown in figure 11 for pressures in the plane of rotation does not seem to occur in the tip Mach number range of the tests where $\\frac{x}{D} \\neq 0$. At all points investigated outside of the plane of rotation the amplitude was found to decrease as the order of the harmonic increased. This is shown in figure 12 where the harmonic amplitude variations for three different tip Mach numbers at several points in the pressure field are given.\n\nComparison of theory with experiment.- In the development of the theory the pressures at a point in space due to the forces distributed over the propeller disk are given by a double integration. The first integration is around the blade path from $\\theta = 0$ to $\\theta = 2\\pi$ and the second integration is along the blade radius from $r = 0$ to $r = R$. For simplification the second integration is eliminated and all forces on the propeller disk are assumed to be concentrated at an effective radius. This effective radius $R_e$ is a function of the blade thrust distribution and torque distribution and the manner in which the forces at each blade element contribute to the free-space pressures at a point in space for a given harmonic. Thus $R_e$ may differ for the various harmonics and may be different for the thrust and torque terms of equation (2).\n\nThe effective radius for a given harmonic was evaluated herein by comparing the calculations with corresponding experimental values. Calculations in figure 13(a) for $R_e = 0.8R$ give good agreement with experiment for the propeller operating at $\\beta_{0.75} = 15^\\circ$ and $M_t = 0.75$. Similar calculations for this propeller at $\\beta_{0.75} = 10^\\circ$ and $M_t = 1.00$ and for $R_e = 0.8R$ overestimate the maximum oscillating pressures. (See fig. 12(b).)\n\nIn figure 14 the experimental and calculated pressures at $\\frac{x}{D} = -\\frac{1}{8}$ are compared for the first three harmonics of the NACA 4-(5)(08)-03 two-blade propeller at $\\beta_{0.75} = 10^\\circ$. The calculated points were obtained by using equation (2) and the force coefficients listed in the figure. Equation (2) predicts pressures over the entire test range of tip Mach numbers with the same amount of accuracy. The deviation then appears to be essentially due to blade loading and not due to tip Mach number. The use of $R_e = 0.8R$ in this case resulted in overestimating all pressures by about 40 percent.", "timestamp": "2026-07-22T06:04:28.079664+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 40, "total_pages": 98, "image_filename": "19930086073_p40.jpg", "text": "```markdown\n38\n\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n-.2\n-.4\n\nLift coefficient, $C_L$\n\n0 0 0 0 0 .1 .2 .3 .4 .5 .6 .7 .8\nDrag coefficient, $C_D$\n\n+45.4 +20.4 0 -10.8 -20.7\nFlap deflection, $\\delta_f$, deg\n\n(b) $C_L$ vs $C_D$.\n\nFigure 7.— Continued.\n\n[Figure: NACA logo]\n\nNACA RM A59E04\n```", "timestamp": "2026-07-22T06:04:30.859048+00:00"}
{"citation_id": "19930082472", "source_url": "https://ntrs.nasa.gov/api/citations/19930082472/downloads/19930082472.pdf", "page_number": 34, "total_pages": 34, "image_filename": "19930082472_p34.jpg", "text": "```markdown\n32\nNACA TN No. 1797\n\n<!-- Image (86, 154, 877, 820) -->\n\nFigure 7. - Characteristics of sections at six spanwise stations of the 45° swept-forward wing.\n```", "timestamp": "2026-07-22T06:04:31.223701+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 99, "total_pages": 99, "image_filename": "19930082511_p99.jpg", "text": "NACA TN No. 1826\n97\n\n[Figure: A diagram showing a complex contour integration path. It features a Cartesian coordinate system with X and Y axes. A semicircular arc in the upper half-plane is shown with arrows indicating a counter-clockwise direction. The radius of the arc is labeled $R \\to \\infty$. The path along the real axis has small semicircular indentations around the points $-h$ and $+h$, with arrows indicating the direction of traversal.]\n\nNACA\n\nFigure 27.-Path of complex contour integration.", "timestamp": "2026-07-22T06:04:35.445091+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930082613_p9.jpg", "text": "8\nNACA TN 1938\n\nThe following observations were made on the type-B specimen subjected to thermal gradients:\n\n1. A thermal gradient produced by heating a louver flap and the surrounding metal to approximately 1700° F for 1 minute caused neither buckling nor cracking.\n\n2. When the area was alternately heated to approximately 1700° F and cooled 30 times, a very slight buckle was formed between the louver and the air-intake holes.\n\n3. When a slightly higher temperature of 1900° F and a more severe temperature gradient were produced in the area immediately upstream and when heating time was increased to 2 minutes, a greater degree of warping occurred. This warpage was not at all comparable in degree or shape to warpage or buckling produced in liners run in accelerated-life determinations or in service.\n\n4. When the louver flaps were heated to approximately 1600° F and the metal downstream was kept cool, the louver flaps were bent outward, approximately halving the gap between the edge of the flap and the liner surface. Alternate heating and cooling moved the louver flap inward and outward.\n\nMetallurgical investigation. - The type-A and type-B liners that failed in service and those that were used for long periods of time in the accelerated-life determinations were severely cracked (table I). Cracks of this type are shown in figures 5 and 6. The longer cracks were chiefly transcrystalline (fig. 7) although most of them had intercrystalline segments. Many branches from the larger cracks were intercrystalline.\n\nBoth surface and subsurface scales were present at most of the edges of larger cracks (figs. 8 to 10). In a few cases, branches from these main cracks appeared to be lined with nothing but surface scale (fig. 11); whereas in many cases branches seemed to be entirely of the subsurface type, (figs. 5 and 9(b)). The separation of layers of scale by metal rather than a crack in some cases leads to the observation that a tubular formation of oxides existed about some of the cracks or branches. Evidence of this type of formation may be seen in figures 5, 8, and 9(b) and a pictorial explanation of what is believed to occur is shown in figure 12. Etching sometimes changes the appearance of the scale until it appears to be a two-phase structure (fig. 10). The innermost oxidized zone turns purple and the outer portion remains a dull gray.", "timestamp": "2026-07-22T06:04:42.374026+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 5, "total_pages": 28, "image_filename": "19930082703_p5.jpg", "text": "NACA TN 1983\n\nconfiguration for the present study. A photograph of the tail-surface installation is given as figure 2, and the location and dimensions are shown in figure 3.\n\nHelicopter C (fig. 4) has the same general arrangement as helicopter A in that a single lifting rotor and torque-counteracting tail rotor are used. It is a two-place helicopter of about 2100 pounds gross weight and has a main-rotor diameter of about 35 feet. This helicopter has a gyroscopic device for improving stability and control characteristics, which for the purposes of this paper may best be viewed as serving to increase the damping moments resulting from angular pitching or rolling velocity of the helicopter. The fuselage configuration of the helicopter was understood to have been found by the manufacturer to result in improved stability characteristics as compared with several alternate fuselage configurations tested.\n\nRESULTS\n\nAll of the pull-up time histories presented start with the helicopter in trim in steady level flight at an indicated airspeed of about 80 miles per hour, which is approximately the cruising speed of these helicopters.\n\nHelicopter A\n\nA time history of a \"pull-and-hold\" maneuver for helicopter A is given in figure 5.\n\nPitching velocity.- The pitching-velocity record shows that maximum angular acceleration is achieved quickly following control displacement, but that little or no tendency to reach a constant value exists, although (aside from the effects of the initially gradual airspeed change), the attainment of a constant angular velocity is basically what is expected from a fixed control displacement.\n\nNormal acceleration.- The normal-acceleration curve appears even more undesirable in nature than the pitching-velocity curve by showing no tendency to reach a constant or maximum value and exhibiting a pause in the development of acceleration following the initial rapid rise. Because pilots have been found to notice rapid changes in normal acceleration of $0.02g$ or even less and because the normal acceleration is the primary measure of the change in flight path being achieved, any illogical development of normal acceleration, particularly a divergent tendency, would be expected to cause adverse pilot impressions.", "timestamp": "2026-07-22T06:04:45.977324+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 3, "total_pages": 26, "image_filename": "19930085485_p3.jpg", "text": "NACA RM No. L8K02 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nAERODYNAMIC CHARACTERISTICS AT SUBSONIC AND TRANSONIC\n\nSPEEDS OF A $42.7^\\circ$ SWEEPBACK WING MODEL HAVING AN\n\nAILERON WITH FINITE TRAILING-EDGE THICKNESS\n\nBy Thomas R. Turner, Vernard E. Lockwood,\nand Raymond D. Vogler\n\nSUMMARY\n\nAn investigation at subsonic and transonic speeds has been performed in the Langley high-speed 7- by 10-foot tunnel to determine the aerodynamic characteristics of a $42.7^\\circ$ sweepback wing with a 20-percent-chord and 50-percent-span outboard aileron. The model had a circular-arc airfoil section and the aileron trailing-edge thickness was modified for the different tests. The investigation was performed in transonic flow over a bump on the tunnel floor and in subsonic flow on one of the tunnel side walls. The Mach number for this investigation varied from about 0.42 to 1.17 and the Reynolds number varied from 800,000 to 1,220,000.\n\nThe data presented indicate that changing the circular-arc aileron contour to a flat-sided aileron contour with finite trailing-edge thickness eliminated reversal of control in most cases and generally improved the aileron control characteristics. The drag coefficient was increased and the aerodynamic center shifted rearward in the subsonic Mach number range.\n\nINTRODUCTION\n\nOne of the many problems arising from the use of high-speed aircraft has been that of securing adequate lateral control, particularly in the transonic speed range. An investigation has been made (reference 1) to determine the aileron control characteristics of a $42.7^\\circ$ sweepback circular-arc wing with various ailerons through a Mach number range from 0.5 to 1.2 by using the transonic bump. The original circular-arc contour aileron gave very low effectiveness in the transonic speed range and the effectiveness reversed for some conditions. An aileron having flat sides and a trailing edge the same thickness as the aileron thickness at the hinge line gave reasonable effectiveness and showed no reversal of effectiveness up to a Mach number of 1.2 (reference 1).\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:04:46.125963+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 11, "total_pages": 44, "image_filename": "19930082566_p11.jpg", "text": "NACA TN No. 1889\n9\n\nequal to approximately zero. Strengths were determined for various\nnumbers of stress application up to about $2 \\times 10^6$ cycles. The S-N\nor $\\sigma$-N diagrams for the four principal stress ratios are shown in\nfigure 12. The data used in plotting the diagrams in figure 13 are\nshown in tables 1 to 4. Since it was necessary to limit the stress\napplications to a low rate to avoid interference of pressure waves,\nthe tests had to be limited to a relatively low number of stress cycles,\nas shown in figure 13.\n\nANALYSIS AND DISCUSSION\n\nThe influence of the principal stress ratio on the fatigue strength\nis shown in figure 14, which shows in a single plot the curves from the\nS-N diagrams of figure 12. The influence of the principal stress ratio\non the fatigue strength can be shown more clearly than in figure 12 by\na comparison of the biaxial fatigue strengths ($\\sigma_1'$ or $\\sigma_2'$) with the\nlongitudinal uniaxial fatigue strengths $\\sigma_{1t}'$. The strength ratios for\nvarious numbers of cycles as shown in figure 13 are in terms of coordi-\nnates $\\sigma_1'/\\sigma_{1t}'$ and $\\sigma_2'/\\sigma_{1t}'$, where $\\sigma_{1t}'$ is the fatigue strength for\nuniaxial longitudinal tension for a given value of N and $\\sigma_1'$\nand $\\sigma_2'$ are the principal stress values for the same value of N.\n\nAttempts were made to compare the test results in figure 15 with\nthe theories of failure (reference 2), but no existing theory was found\nadequate. That is, all the theories require that the material be\nhomogeneous and isotropic, so that the uniaxial strengths $\\sigma_{1t}'$\nand $\\sigma_{2t}'$ in the longitudinal and circumferential directions are equal\naccording to these theories. An examination of figure 13 shows that\nthis is far from being true. Figure 13 indicates that the uniaxial\nfatigue strengths in the circumferential direction may be about 60 per-\ncent of the uniaxial fatigue strengths in the longitudinal direction.\nThat is, the extruded tubular specimens have directional properties\nwith a greater strength in the longitudinal direction. This directional\neffect was also found for the yield and ultimate static strengths in\nreference 1 where the yield strength in the circumferential direction\nwas about 90 percent of that in the longitudinal direction and the\ncorresponding percentage for nominal ultimate strength was about\n80 percent.\n\nTo determine whether a modified maximum-stress theory could be\nused to interpret the foregoing test results if the directional properties\nof the material were considered, the fatigue-strength data from this\nreport and the static-strength data from reference 2 were plotted as shown\nin figures 15 and 16. In plotting figure 16, $\\sigma_{1y}$ and $\\sigma_{2y}$ represent the\nbiaxial yield strengths, and in figure 15, $\\sigma_{1u}$ and $\\sigma_{2u}$ represent the", "timestamp": "2026-07-22T06:04:51.677196+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 11, "total_pages": 50, "image_filename": "19930082592_p11.jpg", "text": "10\nNACA TN 1914\n\nshow an unetched and an etched oxidation interface, respectively, for a 30-percent-tungsten specimen. The unetched field shows that the region of the oxidation moves into the ceramal by following paths into the ceramal rather than moving as a linear front. The etched field clearly shows that the paths being followed by the oxide penetration are the grain boundaries. Mention has previously been made that the grain-boundary phase, which forms when the solubility of the metallic element in the titanium carbide is exceeded, does not occur in the 30-percent-molybdenum ceramal. The 20-percent-tungsten ceramal, however, has a wide grain-boundary phase and the amount of this phase visible in the 10-percent-tungsten ceramal is much less.\n\nThe oxidation penetration along grain boundaries of the 10-percent-tungsten ceramal is comparatively insignificant, as shown by figure 16. When the oxidation proceeds by grain-boundary penetration, the area of oxidation interface is greatly increased and the total oxidation rate is much more rapid. At $2000^\\circ$ F, the oxidation of the titanium carbide is more rapid and this effect coupled with the decreased solubility of oxygen in the WO$_3$ results in a change back from the grain-boundary type of oxidation to a linear oxidation boundary (fig. 17). This change is responsible for the oxidation-rate constants of the ceramals containing high percentages of tungsten at high temperature being of the same order of magnitude. The presence of striations in the oxides of the 5- and 10-percent-tungsten ceramals may also have an effect on the oxidation rate of these ceramals. The layers have a greater resistance to oxygen diffusion while they are building up to a critical thickness than would be encountered if the oxide was torn away from the oxidation interface as soon as it formed, as is the case with the ceramals of higher tungsten content. No attempt was made to determine the relative importance of the three changes in oxidation mechanism described, phase change, striated coating, and grain-boundary oxidation. All three changes would tend to produce the deviations shown in the curves of figure 4(d).\n\nThe oxide coating that forms on the cobalt ceramals is very tightly adherent and dense. In preparing the specimens of cobalt ceramals for examination after oxidation, it was found that the oxide ground to a metallic-like surface, which made it difficult to determine when the scale was completely removed. The oxide appeared to be less amenable to grinding than the ceramal itself because the removed oxides adhered to the diamond grinding wheel. At a magnification of 50 diameters (fig. 18), the coating appears as two separate layers with transition zones at the oxidation interface and between the layers. During the grinding of several specimens, the scale fractured", "timestamp": "2026-07-22T06:04:55.777455+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 76, "total_pages": 96, "image_filename": "19930085880_p76.jpg", "text": "74\nNACA RM No. L9C03\n\n[Figure: A graph plotting Draft, ft (y-axis) against Wetted area, sq ft (x-axis). The y-axis ranges from 0 to .64. The x-axis ranges from 0 to .35. A legend box indicates Speed (fps) with symbols: 10 O, 15 □, 20 ◇, 25 △, 30 ▽. Data points for all speeds are plotted along a single linear trend line rising from the origin. Below the x-axis label is the text \"(e) $\\tau = 20^\\circ$.\"]\n\nWetted area, sq ft\n(e) $\\tau = 20^\\circ$.\nFigure 21.- Concluded.\nNACA", "timestamp": "2026-07-22T06:04:59.559535+00:00"}
{"citation_id": "19930093769", "source_url": "https://ntrs.nasa.gov/api/citations/19930093769/downloads/19930093769.pdf", "page_number": 26, "total_pages": 39, "image_filename": "19930093769_p26.jpg", "text": "NACA RM No. E8L10a CONFIDENTIAL 25\n\n1070\n\nCorrected tail-pipe gas temperature, °R\n\nFuel\nAN-F-58\nGasoline\n\n(e) Altitude, 35,000 feet; flight Mach number, 1.00.\n\nCorrected tail-pipe gas temperature, °R\n\n(f) Altitude, 50,000 feet; flight Mach number, 0.85.\n\nCorrected engine speed, rpm\n\nNACA\n\nFigure 5. - Concluded. Comparison of corrected tail-pipe gas temperature for AN-F-58 fuel and gasoline at various altitudes and flight Mach numbers.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:05:15.473369+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 41, "total_pages": 98, "image_filename": "19930086073_p41.jpg", "text": "```markdown\nNACA RM A9H04\n\nLift coefficient, $C_L$\n\nPitching-moment coefficient, $C_m$\n\nFlap deflection, $\\delta_f$, deg\n\n| Symbol | Value |\n| :--- | :--- |\n| $\\nabla$ | -20.7 |\n| $\\diamond$ | -10.8 |\n| $\\circ$ | 0 |\n| $\\triangle$ | +20.4 |\n| $\\square$ | +45.4 |\n\n(c) $C_L$ vs $C_m$.\n\nFigure 7.- Concluded.\n\n[Figure: A graph plotting Lift coefficient ($C_L$) on the y-axis against Pitching-moment coefficient ($C_m$) on the x-axis. The y-axis ranges from -0.4 to 1.4. The x-axis ranges from 0.12 to -0.28. There are five distinct curves plotted with different symbols corresponding to different flap deflection angles ($\\delta_f$). The curves generally show $C_L$ increasing as $C_m$ becomes more negative, with some curves showing a stall or break at higher lift coefficients. A NACA logo is present in the bottom right corner of the plot area.]\n\n39\n```", "timestamp": "2026-07-22T06:05:17.094202+00:00"}
{"citation_id": "19930082613", "source_url": "https://ntrs.nasa.gov/api/citations/19930082613/downloads/19930082613.pdf", "page_number": 10, "total_pages": 46, "image_filename": "19930082613_p10.jpg", "text": "NACA TN 1938\n9\n\nThe punched edges, as well as all other edges exposed to the gaseous atmosphere, are covered with surface scale; very often sub-surface scale is also present.\n\nThe punched edges of an as-fabricated type-A liner were bordered by a layer of disturbed and elongated grains and were torn in several places (table I and fig. 13). The fissures produced by tearing were large enough to act as stress raisers or potential sources of cracking, the largest being 0.0019 inch long.\n\nIn an ordinary type-B liner run in the accelerated-life run for 16 hours and 40 minutes, some punched edges were found to be lined with small equiaxed grains, less than A.S.T.M. grain size 8, inasmuch as the cold-worked layer formed by the punching operation had recrystallized. Small incipient cracks extended from the punched edges through the layer of fine grains (fig. 14). These incipient cracks were so wide that no definite conclusions could be drawn regarding their origin. It is possible that initial cracking, which extended into the metal deeper than the fine-grained layer, preceded from some of the fissures originally present at the punched edges. Small equiaxed grains were not found on the punched edges of the type-B liners run for 66 hours and 57 minutes in the accelerated-life runs, but distorted grains were present in a few samples. Scaling very probably removed almost all traces of original grain structure at the edges.\n\nQuantitative microscopic comparisons of punched and cracked edges and of inner grains showed no evidence of carburization or decarburization.\n\nYellow intermetallic compounds were found to be scattered uniformly throughout the microstructure and, for this reason, their presence could not be correlated with the failures. Solid nonmetallic inclusions were found in the metal but no abnormal segregations of any type were noticed.\n\nTrial heat treatments. - Heat treatment at 900° F produced no significant difference from the as-fabricated sample (figs. 15(a) and 15(b)). Heating at 1600° F caused precipitation of microconstituents, whereas heating at 2200° F dissolved most of these microconstituents (figs. 15(c) and 15(d)). In addition, heating at 2200° F softened the metal, caused grain growth, and aided recrystallization. No significant differences were observed between samples that were quenched in air or in water (table II).", "timestamp": "2026-07-22T06:05:19.908101+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 2, "total_pages": 47, "image_filename": "19930083221_p2.jpg", "text": "NACA TN No. 1824\n\nTABLE OF CONTENTS\n\nPage\n\nSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:05:20.634022+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930085548_p1.jpg", "text": "RM E8L30\nUNCLASSIFIED\nRESTRICTED\nCOPY NO. 115\nRM No. E8L30\n\nNACA RM No. E8L30\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL STUDY OF LOOP-SCAVENGED COMPRESSION-IGNITION\nCYLINDER FOR GAS-GENERATOR USE\n\nBy Hampton H. Foster, F. Ralph Schuricht\nand Max J. Tauschek\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nTECHNICAL LIBRARY\nAIRESEARCH MANUFACTURING CO.\n9551-9951 SEPULVEDA BLVD.\nINGLEWOOD,\nCALIFORNIA\n\n[Figure: A stamp with text about classified documents and national defense]\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nApril 4, 1949\n\nRESTRICTED\nUNCLASSIFIED", "timestamp": "2026-07-22T06:05:24.245173+00:00"}
{"citation_id": "19930085485", "source_url": "https://ntrs.nasa.gov/api/citations/19930085485/downloads/19930085485.pdf", "page_number": 4, "total_pages": 26, "image_filename": "19930085485_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. L8K02\n\nThe purpose of this investigation was to determine the rolling-moment characteristics of the model of reference 1 with an aileron having flat sides and a trailing-edge thickness less than the thickness at the hinge line, up to a Mach number of 1.17. Lift and drag characteristics were obtained up to a Mach number of 1.15 and the pitching-moment characteristics, up to a Mach number of 0.95.\n\nSYMBOLS AND CORRECTIONS\n\n| | | |\n| :--- | :--- | :--- |\n| $C_L$ | lift coefficient | $\\left(\\frac{L}{\\frac{1}{2}qS}\\right)$ |\n| $C_D$ | drag coefficient | $\\left(\\frac{D}{\\frac{1}{2}qS}\\right)$ |\n| $C_m$ | pitching-moment coefficient | $\\left(\\frac{M'}{\\frac{1}{2}qS\\bar{c}}\\right)$ |\n| $C_{l_a}$ | rolling-moment coefficient produced by aileron | $\\left(\\frac{L'}{qSb}\\right)$ |\n| L | lift, pounds | |\n| D | drag, pounds | |\n| M' | pitching moment about 0.183 mean geometric chord, foot-pounds | |\n| L' | rolling-moment produced by aileron about plane of symmetry, foot-pounds | |\n| S | twice area of semispan model (0.25 sq ft) | |\n| b | twice span of semispan model (1 ft) | |\n| $\\bar{c}$ | wing mean geometric chord (0.259 ft) | |\n| t | ratio of aileron thickness at trailing edge to thickness at hinge line | |\n| $\\alpha$ | angle of attack, degrees | |\n| $\\delta_a$ | aileron deflection, positive when trailing edge is down, degrees | |\n| q | average dynamic pressure over span of model, pounds per square foot $\\left(\\frac{1}{2}\\rho V^2\\right)$ | |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:05:25.123986+00:00"}
{"citation_id": "19930082703", "source_url": "https://ntrs.nasa.gov/api/citations/19930082703/downloads/19930082703.pdf", "page_number": 6, "total_pages": 28, "image_filename": "19930082703_p6.jpg", "text": "4\nNACA TN 1983\n\nPilot's comments.- The pilot's report on this maneuver was that considerable apprehension was felt as a result of the tendency for the helicopter to \"dig in.\" Correspondingly, in normal flying at the same speed, any deviations from steady flight had to be checked at an early stage. The reasons for apprehension with helicopters having these divergent tendencies may be further illustrated by means of figure 6 (for a different helicopter at 65 mph), which is taken from reference 1. The figure shows that, although at the time recovery was initiated the normal acceleration differed from that for level flight by only -0.3g, during recovery an increment of 0.8g occurred even though the control stick was full forward by the time this increment was reached. The maneuver was checked at this point only with the aid of other flight controls. Miscellaneous measurements obtained with helicopter A indicate that at 80 miles per hour it would exhibit characteristics generally similar to, though somewhat milder than, those shown in figure 6.\n\nIn addition to the divergent tendency, the pilot reported difficulty in anticipating, during the first 1 or 2 seconds, the rapidity with which the divergence would later take place. In pull-ups started at lower speeds, a maximum acceleration value could be reached, but a similar difficulty in anticipating the eventual result was noted.\n\nStick-fixed oscillations.- Following a nose-up disturbance at 80 miles per hour, it was necessary to effect recovery during the first nose-down motion. In other words, only a part of an oscillation could be tolerated.\n\nHelicopter B\n\nA time history of a pull-and-hold maneuver for helicopter B is given in figure 7.\n\nPitching velocity.- The maximum angular acceleration is again reached quickly following control displacement and, in this case, a tendency to reach a constant (or at least a maximum) angular velocity is almost immediately evident; that is, the curve of angular velocity is definitely concave downward. Maximum angular velocity is reached in about $1\\frac{1}{2}$ seconds, which it is understood would not be objectionable for airplanes.\n\nNormal acceleration.- The normal-acceleration curve again exhibits the pause in development (following the initial rapid rise) as noted for helicopter A, but by the end of about 2 seconds (at which time recovery was applied with helicopter A), a tendency to reach a constant or maximum value was evident and the pilot held the deflection until after", "timestamp": "2026-07-22T06:05:25.323147+00:00"}
{"citation_id": "19930086097", "source_url": "https://ntrs.nasa.gov/api/citations/19930086097/downloads/19930086097.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930086097_p32.jpg", "text": "30\nCONFIDENTIAL\nNACA RM A9H11\n\n<!-- Image (144, 99, 817, 472) -->\n\nFigure 11. -Profile drag of wings 1, 2, 3, and 4; $M_\\infty=2.0$, artificial roughness added.\n\n<!-- Image (175, 527, 812, 901) -->\n\nFigure 12. -Lift curves of wings 5, 6, and 7; $M_\\infty=1.5$, $Re=0.5 \\times 10^6$.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:05:29.551691+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 1, "total_pages": 17, "image_filename": "19930085572_p1.jpg", "text": "RESTRICTED\nCOPY NO.\nRM No. E8L02\n\nNACA RM No. E8L02\n\nNACA\n\nRESEARCH MEMORANDUM\n\nCOMPARISON OF FLIGHT PERFORMANCE OF AN-F-58\nAND AN-F-32 FUELS IN J35 TURBOJET ENGINE\n\nBy Loren W. Acker and Kenneth S. Kleinknecht\n\nLewis Flight Propulsion Laboratory\nCleveland, Ohio\n\nCLASSIFICATION CHANGED TO\nUNCLASSIFIED\nAUTHORITY CROWLEY CHANGE #2034\nDATE 12-14-53\nT.C.F.\n\nCLASSIFIED DOCUMENT\n\nThis document contains classified information affecting the National Defense of the United States within the meaning of the Espionage Act, USC 50:31 and 32. Its transmission or the revelation of its contents in any manner to an unauthorized person is prohibited by law. Information so classified may be imparted only to persons in the military and naval services of the United States, appropriate civilian officers and employees of the Federal Government who have a legitimate interest therein, and to United States citizens of proven loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nApril 7, 1949\n\nRESTRICTED", "timestamp": "2026-07-22T06:05:34.503683+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 12, "total_pages": 44, "image_filename": "19930082566_p12.jpg", "text": "```markdown\n10\nNACA TN No. 1889\n\nbiaxial nominal ultimate strengths. If a modified maximum-stress theory\nis to agree with the test results in figure 16, the ratios $\\sigma_1'/\\sigma_{1y}$\nand $\\sigma_2'/\\sigma_{2y}$ should remain constant for all values of the principal\nstress ratio and for each value of N. Figure 16 shows that the stress\nratios $\\sigma_1'/\\sigma_{1y}$ and $\\sigma_2'/\\sigma_{2y}$ are not constant. Figure 15 shows that\nthere is also a variation in the strength ratios $\\sigma_1'/\\sigma_{1u}$ and $\\sigma_2'/\\sigma_{2u}$\nwith variation in the principal stress ratio. That is, a modified\nmaximum-stress theory based on either yield or ultimate strengths does\nnot agree with test results. However, figures 15 and 16 are of\nvalue in showing the relation between the fatigue and static strengths\nof the material for various ratios of the principal stresses.\n\nIn figure 17 a comparison is made between the S-N diagram based on\nthe longitudinal fatigue-stress results reported in the foregoing\nparagraphs for $\\sigma_2/\\sigma_1 = 0$ and the S-N diagram based on data given in\nreference 3 for 0.2-inch-diameter specimens. Figure 17 shows that there\nis an appreciable reduction in fatigue strengths for the tubular specimens,\nsince the S-N curve for these specimens lies well below that for the solid\nspecimens.\n\nFigure 18 is a photograph of typical fractured specimens. For stress\nratios of $\\sigma_2/\\sigma_1 = 0.5$, the specimens fractured circumferentially. The\nplane of the fracture was at an angle of about $45^\\circ$ to the surface of the\ntube. For stress ratios $\\sigma_2/\\sigma_1 = 1.0$ and $\\sigma_2/\\sigma_1 = 2.0$, failure was\nproduced by small cracks in the longitudinal direction about 1/2\nto 1 inch in length.\n\nCONCLUSIONS\n\nBiaxial tensile fatigue strengths of 24S-T aluminum-alloy tubing\nwere obtained for various ratios of the biaxial maximum stresses and\nwith the minimum stresses approximately equal to zero. The test results\nshow that uniaxial fatigue-strength values in the longitudinal direction\ncannot be used to predict the fatigue strength, and that the biaxial\nfatigue strength may be as low as 50 percent of the uniaxial fatigue\nstrength.\n\nThe Pennsylvania State College\nState College, Pa., December 13, 1947\n```", "timestamp": "2026-07-22T06:05:36.594030+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 25, "total_pages": 78, "image_filename": "19930082483_p25.jpg", "text": "NACA TN No. 1807\n23\n\ncalibration of the lubricating system mock-up. The temperature\nrise of the oil was obtained from thermocouples in the oil lines\nbefore and after each bearing. The specific heat of the oil was\nbased on the mean oil temperature in the bearing.\n\nPumping loss. - This partial-admission loss was calculated\nfrom equation (21), using an experimentally determined\n\n$$\\lambda = 15.19$$\n\nand the equation then corrected to sea level to give\n\n$$\\text{pumping loss} = \\frac{15.19 \\left(\\frac{l}{D}\\right) D^5 \\left(\\frac{N}{1000}\\right)^3 (1-F) \\rho_d}{\\delta_i \\sqrt{\\theta_i}} \\quad (40)$$\n\nThis equation, which has the advantage of dimensional correct-\nness, may be adjusted for a particular application merely by\nselection of the term $\\lambda$. In general,\n\n$$\\lambda = 27.99 \\text{ for uncovered inactive blading}$$\n$$= 12.93 \\text{ for inactive blading having a cover}$$\n$$\\text{with a 3/16-inch clearance}$$\n\nThe configuration investigated does not correspond to either\nof the preceding conditions for which a numerical value for $\\lambda$ is\nlisted, but is somewhere between the two. In order to determine\na more appropriate value for the constant, the smaller turbine,\npreviously mentioned, which had a configuration substantially\nsimilar to that shown in figure 7, was motored and measurements of\nthe pumping loss were obtained. The pumping-loss expression was\nrearranged to make $\\lambda$ the dependent variable, and values of $\\lambda$\nwere obtained over a wide range of rotor speeds and gas densities.\n\nLittle variation in the value of the coefficient was encoun-\ntered and a mean value of 15.19 was obtained for $\\lambda$.\n\nDriving-fluid losses. - Because of the difficulties involved\nin obtaining gas-state measurements to evaluate driving-fluid\nlosses, these losses were experimentally determined for this tur-\nbine and compared with the general expression, equation (27), to\nverify that the losses do vary as $K_{III} \\bar{\\rho}_i N$.\n\nThe driving-fluid losses are experimentally evaluated at\n$120^\\circ$ and $180^\\circ$ admission, using equation (26); the blade power was", "timestamp": "2026-07-22T06:05:40.704482+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 4, "total_pages": 62, "image_filename": "19930082918_p4.jpg", "text": "NACA TN 1940\n3\n\nThis report presents the results obtained to date on bar-stock material, solution-treated at $2200^\\circ$ F for 10 hours and aged for time periods up to 1000 hours at $1200^\\circ$, $1400^\\circ$, and $1600^\\circ$ F. The creep and rupture data were confined to short time periods at $1200^\\circ$ F under a restricted stress range. This is, therefore, in the nature of a report on the techniques developed and progress made to date. A large amount of additional experimental work must be done before all temperatures, stresses, and time periods of interest are related to the structural conditions of the alloy resulting from the wide range of possible prior treatments.\n\nThe investigation is part of a research program on heat-resistant alloys for aircraft propulsion systems conducted at the Engineering Research Institute of the University of Michigan under the sponsorship and with the financial assistance of the National Advisory Committee for Aeronautics.\n\nTEST MATERIALS\n\nLow-carbon N-155 alloy bar stock was used in this investigation. It represents part of the complete product of one ingot of heat A-1726 rolled into the 7/8-inch broken-corner square stock. Each bar from the ingot was numbered so that its position relative to the original ingot was known. The particular material considered herein came from the center of the ingot.\n\nThe composition of heat A-1726 reported by the supplier is listed as follows, together with the results of the analysis by the University of the bar from the center of the particular ingot being considered:\n\n| | Chemical composition (percent) | | | | | | | | | |\n| :--- | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: | :---: |\n| | C | Mn | Si | Cr | Ni | Co | Mo | W | Cb | N |\n| Supplier's heat analysis | 0.13 | 1.63 | 0.42 | 21.22 | 19.00 | 19.70 | 2.90 | 2.61 | 0.84 | 0.13 |\n| Bar from center of ingot (Univ. of Mich.) | .13 | 1.43 | .34 | 20.73 | 18.92 | 19.65 | 3.05 | 1.98 | .98 | .14 |\n\nThe appendix gives the complete processing schedule reported for the ingot of heat A-1726 from which the test stock was taken.", "timestamp": "2026-07-22T06:05:42.379521+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 2, "total_pages": 149, "image_filename": "19930083192_p2.jpg", "text": "```markdown\n# TABLE OF CONTENTS\n\nPage\n\nSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 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. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .", "timestamp": "2026-07-22T06:05:43.852745+00:00"}
{"citation_id": "19930082592", "source_url": "https://ntrs.nasa.gov/api/citations/19930082592/downloads/19930082592.pdf", "page_number": 12, "total_pages": 50, "image_filename": "19930082592_p12.jpg", "text": "NACA TN 1914\n\nalong the transition zone between the two layers and the inner layer was observed to have a characteristic green color. This inner layer appeared to be hard, dense, and tightly adhering to the ceramal. No evidence of separation at the oxidation interface was found. In regard to deterioration or loss of cross-sectional area by oxidation, this inner oxide layer can be considered an integral part of the ceramal. As previously mentioned, the coatings on the molybdenum and tungsten ceramals are not tightly adherent and are readily removable.\n\nNo significant change in the oxidation-rate constants of the cobalt ceramals is evident at the temperatures considered, excepting the 30-percent-cobalt ceramal at $1625^\\circ$ F, which has a value higher than any of the other temperature-composition combinations (fig. 7). Metallographic study of these coatings did not reveal any explanation for the abnormality of the 30-percent-cobalt ceramal. None of these ceramals show any sign of intergranular oxide penetration and in all cases the scales appear similar. The photomicrograph (fig. 19) shows that the oxidation mechanism of cobalt ceramals is not as simple as that of the tungsten and molybdenum ceramals. At the oxidation interface, a thin layer of material forms that is probably a transition zone containing the partly oxidized ceramal material and the oxidation products first formed. Near this layer, the scale is a two-phase material with light insular particles in a darker matrix. Farther from the oxidation interface, the light phase rather suddenly disappears leaving only the dark phase near the outer surface. Examination by X-ray diffraction indicated the presence of $\\text{CoO}\\cdot\\text{Co}_2\\text{O}_3$ close to the oxidation interface and $\\text{CoTiO}_3$ at the outer face of the inner layer of scale. When the oxide chipped off during grinding, the fracture occurred just inward from the area in which the last of the light phase finally disappears. Only $\\text{TiO}_2$ was found to be definitely present in the outer dark layer. Apparently the $\\text{CoO}\\cdot\\text{Co}_2\\text{O}_3$ and a portion of the $\\text{TiO}_2$ that forms at the oxidation interface combine to form $\\text{CoTiO}_3$ and the $\\text{CoTiO}_3$ enters into solution in the $\\text{TiO}_2$. The behavior during oxidation of the 30-percent-cobalt material is probably the result of the complex reactions taking place throughout the oxygen-diffusion zone and of the effect variations in temperature and concentration of cobalt may have on these reactions.\n\nOf the compositions considered, the use of cobalt, tungsten, and molybdenum as metallic additions will result in a marked difference in oxidation resistance. With regard to oxidation, the tungsten compositions are superior to the molybdenum compositions and the cobalt compositions are in general superior to both.", "timestamp": "2026-07-22T06:05:47.152463+00:00"}
{"citation_id": "19930085572", "source_url": "https://ntrs.nasa.gov/api/citations/19930085572/downloads/19930085572.pdf", "page_number": 2, "total_pages": 17, "image_filename": "19930085572_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:06:04.911597+00:00"}
{"citation_id": "19930085880", "source_url": "https://ntrs.nasa.gov/api/citations/19930085880/downloads/19930085880.pdf", "page_number": 77, "total_pages": 96, "image_filename": "19930085880_p77.jpg", "text": "NACA RM No. L9C03\n75\n\n[Figure: A line graph plotting Load (lb) against Wetted area (sq ft). The graph contains five data series representing different speeds (10, 15, 20, 25, 30 fps). An inset diagram in the upper left corner shows a schematic of a hull cross-section.]\n\nLoad, lb\n32\n28\n24\n20\n16\n12\n8\n4\n0\n\nSpeed (fps)\n30\n25\n20\n15\n10\n\n0 .05 .10 .15 .20 .25 .30 .35\nWetted area, sq ft\n\nNACA\n\n(a) $\\tau = 40$.\n\nFigure 22.- Variation of load with wetted area. Model 250D.", "timestamp": "2026-07-22T06:06:06.376887+00:00"}
{"citation_id": "19930082566", "source_url": "https://ntrs.nasa.gov/api/citations/19930082566/downloads/19930082566.pdf", "page_number": 13, "total_pages": 44, "image_filename": "19930082566_p13.jpg", "text": "NACA TN No. 1889\n\n11\n\nREFERENCES\n\n1. Marin, Joseph: Interpretation of Experiments on Fatigue Strength of Metals Subjected to Combined Stresses. The Welding Jour., vol. VII, no. 5, May 1942, pp. 245-s – 248-s. (See also, Marin, J.: Mechanical Properties of Materials and Design. McGraw-Hill Book Co., Inc., 1942, ch. 4.)\n\n2. Marin, Joseph, Faupel, J. H., Dutton, V. L., and Brossman, M. W.: Biaxial Plastic Stress-Strain Relations for 24S-T Aluminum Alloy. NACA TN No. 1536, 1948.\n\n3. Anon: Alcoa Structural Handbook. ALCOA, 1945.", "timestamp": "2026-07-22T06:06:09.935503+00:00"}