PDF/ocr_dataset_1077.jsonl

{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 53, "total_pages": 60, "image_filename": "19930085862_p53.jpg", "text": "NACA RM No. L9A07\n51\n\n<!-- Image (231, 109, 852, 846) -->\n\n(b) $C_L$, $C_D$, and $C_m$ against $\\alpha$.\nFigure 19.— Concluded.", "timestamp": "2026-07-22T06:49:15.787378+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 39, "total_pages": 92, "image_filename": "19930085870_p39.jpg", "text": "```markdown\n40\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n<!-- Image (65, 110, 931, 904) -->\n\n(j) Wing 10. w=1066; R = 740,000.\nFigure 5. - Continued.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:49:16.399872+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 7, "total_pages": 30, "image_filename": "19930085975_p7.jpg", "text": "NACA RM L9E10 CONFIDENTIAL 5\n\nStatic loading of the ailerons and calculations of the twist of the wing indicated that errors arising from points (1) and (2) are negligible; however, calculations of the maximum change in angle of attack of the wing tip of the sweptback wings due to bending of the wing panel indicate an appreciable change in the angle. This change in angle of attack is in such a direction as to reduce the rolling moment which is being produced by the ailerons and is only important when the model is restrained in roll for the static tests. When the model has attained a steady rate of roll, in the free-roll tests, the damping moment of the wing balances the aileron rolling moment. In this condition the lateral center of pressure for the damping moment and the aileron load are at slightly different spanwise locations, and there is some slight distortion of the wing. This distortion is negligible, however, when compared with the distortions in the restrained condition and the wing can be considered essentially rigid during the free-roll tests.\n\nThe rolling moment $L_{\\text{test}}$ which is measured during the static test is\n\n$$\nL_{\\text{test}} = L_{\\delta} - L_{\\epsilon}\n\\tag{1}\n$$\n\nwhere $L_{\\epsilon}$ is the rolling moment lost due to the bending of the wing. This increment of rolling moment $L_{\\epsilon}$ can be estimated by the relation\n\n$$\nL_{\\epsilon} = C_{l_p} \\epsilon q S b\n\\tag{2}\n$$\n\nwhere $C_{l_p}$ is the damping coefficient for the rigid wing. This estimation involves the assumption that the angle of attack due to distortion varies linearly from zero at the root to $\\epsilon$ at the tip. This is not strictly correct of course, but the assumption is believed to give a good first approximation of the increment of rolling moment lost due to bending.\n\nIn order to determine $\\epsilon$, the applied rolling moments $L_{\\text{test}}$ were approximated by concentrated loads applied to the wing at the center of load calculated on the basis of unswept-wing theory (reference 6). The change in angle of attack at the wing tip $\\epsilon$ was measured relative to the root chord and the rate of change with rolling moment $\\Delta \\epsilon / \\Delta L_{\\text{test}}$ was determined.\n\nEquation (2) can be written\n\n$$\nL_{\\epsilon} = C_{l_p} \\frac{\\Delta \\epsilon}{\\Delta L_{\\text{test}}} L_{\\text{test}} q S b\n\\tag{3}\n$$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:17.201585+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 19, "total_pages": 22, "image_filename": "19930085928_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:49:18.011895+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 18, "total_pages": 34, "image_filename": "19930085913_p18.jpg", "text": "```markdown\nTABLE I.- EXPERIMENTAL DATA - Continued\n\n| Model | Run | $q_r$ (lb/sq ft) | $V_r$ (fps) | Mach number | Distance of weight from root (percent l) | Frequencies (cps) | | | Phase-angle relationship of bending and torsional stresses. (Ref indicates reference strain-gage trace) | | | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | **Natural** | | | **Flutter** | **2nd natural mode** | | | | **3rd natural mode** | | | | **Flutter mode** | | | |\n| | | | | | | 1st | 2nd | 3rd | | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) | 1 | 2 (deg) | 3 (deg) | 4 (deg) |\n| Model C<br>Swept untapered wing;<br>$\\Lambda = 60^\\circ$<br>Weight moved along leading<br>edge; $e_w = -1$<br>Reynolds number $\\cong 7306.2V_r$ | 64 | 40.92 | 192.7 | 0.1670 | -19.44 | 2.94 | 17.52 | 22.35 | 13.33 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Ref | 180 | 14 | 24 |\n| | 65 | 40.72 | 193.1 | .1670 | 0 | 2.93 | 17.55 | 21.54 | 13.62 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Ref | 180 | 14 | 19 |\n| | 66 | 33.50 | 174.6 | .1510 | 13.88 | 2.92 | 13.70 | 19.7 | 11.41 | Ref | 0 | 0 | 0 | Ref | 0 | 0 | --- | Ref | --- | 0 | 0 |\n| | 67 | 30.98 | 168.2 | .1450 | 19.44 | 2.92 | 11.43 | ----- | 10.19 | Ref | 0 | 0 | 0 | | No record | | | Ref | 36 | 14 | 18 |\n| | 68 | 25.94 | 153.8 | .1325 | 25.00 | 2.88 | 9.71 | 20.51 | 8.75 | Ref | --- | - | 0 | --- | Ref | - | 180 | Ref | 55 | 12 | 30 |\n| | 69 | 27.78 | 159.2 | .1373 | 30.55 | 2.88 | 8.40 | 20.83 | 8.12 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Ref | 64 | 33 | 39 |\n| | 70 | 36.84 | 183.3 | .1585 | 36.11 | 2.80 | 7.74 | 21.21 | 7.01 | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | Ref | 105 | 105 | 79 |\n| | 71 | 96.88 | 299.3 | .2600 | 41.67 | 2.70 | 7.32 | 20.83 | 6.79<br>$^a$30.40 | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | Refer to record | | | |\n| | 72 | 126.10 | 343.3 | .2989 | 47.22 | 2.53 | 7.35 | 20.51 | $^a$34.20 | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | Refer to record | | | |\n| | 73 | 142.70 | 366.1 | .3189 | 52.77 | 2.38 | 7.59 | 20.33 | $^a$37.90 | Ref | 180 | 0 | 0 | Ref | 0 | 0 | 180 | Refer to record | | | |\n| | 74 | 222.00 | 465.8 | .4050 | 58.33 | 2.13 | 8.10 | 20.00 | 6.50<br>$^a$40.00 | Ref | 180 | - | 0 | Ref | 0 | 0 | 180 | Refer to record | | | |\n| | 75 | Limiting tunnel velocity | | | 63.89 | 2.00 | 8.82 | 19.23 | ----- | Ref | --- | - | 0 | Ref | 0 | 0 | 180 | No flutter | | | |\n| | 76 | 216.60 | 459.8 | .3995 | 69.44 | 1.81 | 9.46 | 19.40 | $^a$22.20 | Ref | --- | 0 | 0 | --- | Ref | 0 | 180 | Ref | 79 | 324 | 292 |\n| | 77 | 126.30 | 345.4 | .2996 | 75.00 | 1.68 | 10.20 | 18.75 | 18.33 | Ref | 0 | 0 | 0 | --- | Ref | - | 180 | Ref | 129 | 0 | 324 |\n| | 78 | 68.53 | 251.2 | .2174 | 80.56 | 1.59 | 11.11 | 17.78 | 17.02 | Ref | 0 | 0 | 0 | --- | Ref | - | 180 | Ref | 162 | --- | 0 |\n| | 79 | 36.75 | 183.4 | .1581 | 86.11 | 1.45 | 11.76 | 16.95 | 16.00 | Ref | 0 | 0 | 0 | Ref | 180 | 180 | 0 | Ref | 152 | 158 | 347 |\n| | 80 | 21.11 | 138.8 | .1195 | 91.67 | 1.35 | 11.70 | 15.91 | 14.67 | Ref | 0 | 0 | 0 | Ref | 180 | --- | 0 | Ref | 165 | --- | 350 |\n| | 81 | 15.28 | 118.1 | .1015 | 97.22 | 1.24 | 11.32 | 15.10 | $^a$13.33 | Ref | 0 | 0 | 180 | Ref | 180 | 0 | 0 | Ref | 125 | 38 | 346 |\n\n[Figure: Diagram of wing model showing sweep angle $\\Lambda$, length $l$, chord $b$, and weight position $2b$]\n\n$^a$Note oscillograph record, figure 1.\n\nNACA\n17\n```", "timestamp": "2026-07-22T06:49:20.317884+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 5, "total_pages": 46, "image_filename": "19930085983_p5.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 3\n\n$$C_{m_{\\delta}}^* = \\left[ \\frac{(C_m)_{\\delta=-4^\\circ} - (C_m)_{\\delta=0^\\circ}}{-4} \\right]$$\n\n| Symbol | Definition |\n| :--- | :--- |\n| H | hinge moment, foot-pounds |\n| M | Mach number $\\left( \\frac{V}{a} \\right)$ |\n| n | normal acceleration factor |\n| p | angular velocity in roll, radians per second |\n| q | dynamic pressure $\\left( \\frac{1}{2}\\rho V^2 \\right)$, pounds per square foot |\n| R | Reynolds number $\\left( \\frac{\\rho V \\bar{c}}{\\mu} \\right)$ |\n| S | wing area, square feet |\n| V | free-stream velocity, feet per second |\n| $V_v$ | sinking speed, feet per second |\n| $V_G$ | gliding speed, miles per hour |\n| y | lateral ordinate, feet |\n| $\\alpha$ | angle of attack of root chord line, degrees |\n| $\\alpha_t$ | angle of twist with reference to root chord (positive for washin), degrees |\n| $\\alpha_u$ | angle of attack of root chord line, uncorrected for tunnel-wall interference, degrees |\n| $\\delta$ | elevon deflection measured in planes perpendicular to the elevon hinge axes (positive downward), degrees |\n| $\\delta_u$ | elevon deflection uncorrected for angular distortion due to load, degrees |\n| $\\delta_{Lu}$ | left elevon deflection uncorrected for angular distortion due to load, degrees |\n| $\\delta_{Ru}$ | right elevon deflection uncorrected for angular distortion due to load, degrees |\n| $\\delta_T$ | arithmetic sum of positive and negative elevon deflections, degrees |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:24.941214+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 2, "total_pages": 132, "image_filename": "19930085990_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:49:25.967663+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 61, "total_pages": 62, "image_filename": "19930082918_p61.jpg", "text": "72\nNACA TN 1940\n\nMaximum true strain, in./in.\n.25\n.20\n.15\n.10\n.05\n0\n\n100-hour age\n1-hour age\n1/2-hour age\nUnaged\n10-hour age\n\n.1\n1.0\n10\n100\n\n1200° F rupture time, hr\n\nFigure 20.- Maximum true strain at rupture of low-carbon M-155 alloy aged at 1600° F, solution-treated 10 hours at 2200° F, and water-quenched.\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T06:49:29.977651+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 50, "total_pages": 149, "image_filename": "19930083192_p50.jpg", "text": "46 NACA TN 1976\n\nFor gradient distances other than zero or for those cases in which the pitch is a factor, the discrepancies between calculation and experiment become larger as the gradient distance increases until, as indicated by figure 35, the errors are serious for gradient distances of 16 to 20 chords. In the case of both the conventional airplanes (fig. 35) and the canard airplane (fig. 38(b)), the calculations for the longer gust-gradient distance are unconservative and lead to smaller predicted load factors. In the case of the tailless airplane (fig. 38(a)) the calculated values are in good agreement. Consideration of the data presented in figure 32, in which the effect of pitch has been neglected in the calculations, indicates that the neglect of pitch for the longer gust-gradient distances will be conservative. The good agreement for the tailless airplane (fig. 38(a)) at all gradient distances compared to the variation in agreement between calculation and experiment for the other cases indicates that the introduction of the tail surface and possibly the fuselage has much to do with the adequacy of the detailed calculations of airplane response to a gust. The errors indicated for the longer gust-gradient distances as compared to the sharper gust indicate that the pitch effect is probably the least accurately predicted quantity and, therefore, leads to the errors noted.\n\nFrom the results obtained it appears that the wing load due to a gust can be calculated within 10 percent for gradient distances up to 10 chords. The accuracy in any individual case or for large gradient distances cannot be estimated because of pitch and secondary effects that may be significant but are not recognized until test results are available. Minor changes in variables such as downwash can lead to serious variation of numerical results, and the estimation of loads by detailed or extensive calculations is not of sufficient accuracy to warrant much confidence in the results.\n\nELASTIC-AIRPLANE REACTIONS\n\nSince the loads imposed when airplanes encounter gusts are applied suddenly, the dynamic response of the airplane structure has been of concern since the initiation of gust-load studies. A number of studies (references 2 and 35 to 40) have been made at various times to evaluate the importance of dynamic response and of the various parameters involved. In 1939, projected airplane designs indicated that dynamic response might be of concern and an analytical and experimental study was undertaken. The results of this investigation are presented in reference 38.\n\nIn this section the results obtained in reference 38 are summarized as well as some results from unpublished studies. The significance of dynamic response and the importance of the various parameters in producing dynamic responses in airplane structures are of chief concern.", "timestamp": "2026-07-22T06:49:30.171636+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 54, "total_pages": 72, "image_filename": "19930085491_p54.jpg", "text": "```markdown\nNACA RM No. A8J04\n\n$M_0 = 1.53$\n$\\Lambda_{LE} = 67.0^\\circ$\n— Approximate linear theory\n$\\circ$ $R=0.62$ million\n$\\diamond$ $R=0.95$ million\n\nCONFIDENTIAL\n\n<!-- Image (116, 125, 402, 784) -->\n\n<!-- Image (432, 125, 840, 784) -->\n\n(d) WF-67\n\nFigure 7- Continued.\n\nCONFIDENTIAL\n\nWF-67\n\n53\n```", "timestamp": "2026-07-22T06:49:37.719518+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 78, "total_pages": 78, "image_filename": "19930082483_p78.jpg", "text": "```markdown\n76\n\n200\nAdmission\n(deg)\n180\n160\nPower corrected to sea-level, hp\n120\n120\n80\nPower\nGross\nBlade\nNACA\n40\n2000\n4000\n6000\n8000\n10,000\n12,000\n14,000\nCorrected rotor speed, $N/\\sqrt{\\theta_i}$, rpm\n\nFigure 14. - Variation of driving-fluid losses (shaded areas\nobtained as difference between gross power and blade power)\nwith rotor speed for 180° and 120° admission at total-pressure\nratio of 2.0.\n\nNACA TN No. 1807\n1032\n```", "timestamp": "2026-07-22T06:49:38.382353+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 86, "total_pages": 98, "image_filename": "19930086073_p86.jpg", "text": "84\n\nLift coefficient, $C_L$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T06:49:41.581218+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 26, "total_pages": 52, "image_filename": "19930085911_p26.jpg", "text": "1152\n\nNACA RM E9F22\n\nCONFIDENTIAL\n\nStation 0 1 2 3 4 5 6 7\n26\" 87.50\" 8\" 34.56\"\n10\" 12.25\" 50° 8\" 15.94\" 11.15\"\n3.65\" 3.75\" 56\" 22\"\nPitot tube and telemetering antenna\nTelemetering equipment\nHelium coils\nFuel-spray ring\nFuel tank\nFlares\nDucted-type flame holder\nTail fins\n\nFigure 3. - Schematic cross-sectional diagram of supersonic 16-inch ram-jet unit. (Dimensions given for model A.)\n\nNACA\n\nCONFIDENTIAL\n\n25", "timestamp": "2026-07-22T06:49:44.536118+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 18, "total_pages": 47, "image_filename": "19930085919_p18.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 17\n\n[Figure: A man in a suit is crouching beside a large, swept-back wing model mounted on a platform inside a wind tunnel. The wing is angled upward and extends diagonally across the frame. In the background, there are windows or viewing ports on the wall. The NACA logo and identifier “A-12315” appear in the lower right corner of the image.]\n\nFigure 1.— The wing mounted in one of the Ames 7- by 10-foot wind tunnels.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:49:49.723717+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 4, "total_pages": 25, "image_filename": "19930085979_p4.jpg", "text": "```markdown\nNACA RM E5E12\n3\n\nTemperatures on the surface of the accessory housing were measured by 17 flush-type skin thermocouples.\n\nThe state of the gas in the plenum chamber was measured by four thermocouple probes located $90^\\circ$ apart in the plane of the orifices and by four static-pressure taps in the rear wall of the plenum chamber.\n\nThe inlet-lip pressure distribution was measured by means of pressure belting cemented to the lip surface. Lip-surface temperatures were measured by thermocouples welded to the inlet lip.\n\nMass flow through the model and inlet-velocity ratio were controlled by an electrically driven tail cone (fig. 2).\n\nPROCEDURE\n\nAerodynamic investigation without bleedback. - An aerodynamic investigation of the model with orifices was conducted to determine mass-flow characteristics, lip-pressure distribution, and ram-pressure recovery as a function of inlet-velocity ratio and angle of attack. Tunnel-air velocities ranged from approximately 200 to 460 feet per second. At each tunnel-air velocity the angle of attack was varied from $0^\\circ$ to $8^\\circ$ and for each angle of attack the inlet-velocity ratio was varied from 0.68 to 0.89.\n\nAerodynamic investigation with cold-gas bleedback. - An aerodynamic investigation of the model was conducted to determine the effect of cold-gas bleedback on mass-flow characteristics, lip-pressure distribution, and ram-pressure recovery. This investigation was conducted at an angle of attack of $0^\\circ$ at a fixed tail-cone position corresponding to an inlet-velocity ratio of 0.89 without bleedback. Tunnel-air velocities ranged from 200 to 460 feet per second and bleedbacks ranged from 2.5 to 9.8 percent.\n\nAerodynamic investigation with hot-gas bleedback. - Several orifice configurations were investigated in order to obtain a configuration that would give the most uniform temperature distribution inside the model for a range of values of tunnel-air velocity, angle of attack, gas flow, and gas temperature. The configuration selected consisted of twelve 17/32-inch-diameter orifices equally spaced around the inlet.\n\nThe effect of hot-gas bleedback on the temperature distribution inside the model, mass-flow characteristics, and ram-pressure recovery for the optimum orifice configuration was determined as a\n```", "timestamp": "2026-07-22T06:49:56.394191+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 85, "total_pages": 114, "image_filename": "19930086061_p85.jpg", "text": "NACA RM L9J07\n81\n\n<!-- Image (267, 100, 729, 650) -->\n\n(d) Three-dimensional top view normal to\nplane of chord lines; $\\alpha=20^\\circ$; $\\psi=20^\\circ$.\n\n<!-- Image (202, 719, 789, 909) -->\n\n(e) Three-dimensional rear view parallel\nto air stream; $\\alpha=20^\\circ$; $\\psi=20^\\circ$.\n\nFigure 30.- Concluded.", "timestamp": "2026-07-22T06:49:57.126579+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 54, "total_pages": 60, "image_filename": "19930085862_p54.jpg", "text": "52\nNACA RM No. L9A07\n\n.01\n0\n$C_n$\n-.01\n-.02\n\nSpoiler Location of\nSpan outboard end\n$\\diamond$ 0.775b/2 0.975b/2\n$\\diamond$ .375b/2 .975b/2\n$\\square$ .40b/2 .60b/2\n\n0\n-.01\n-.02\n$C_l$\n-.03\n-.04\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n[Figure: NACA logo]\n\nFigure 20.— Effects of 0.10c projection step spoilers on characteristics\nof wing with extensible leading-edge flaps and trailing-edge split\nflaps.", "timestamp": "2026-07-22T06:49:57.886702+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 8, "total_pages": 30, "image_filename": "19930085970_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM A9E09\n\nMinimum drag coefficient.- The variation of the minimum drag coefficient of the model with Mach number is shown in figure 8. It can be seen that no appreciable change in minimum drag coefficient occurred at subsonic Mach numbers although there was a slight rise in the values between 0.90 and 0.95 Mach number. This variation is also indicated by the results of reference 3 obtained at a higher Reynolds number. Values of minimum drag coefficient calculated by the methods of reference 8 are shown for both the fully laminar and fully turbulent boundary-layer conditions, since it was not possible to assess the exact proportions of each type of flow that existed on the model. It is seen that the observed and calculated trends of the minimum drag coefficient agree well at subsonic Mach numbers and the measured values fall well within the indicated skin friction limits. The observed increase in minimum drag coefficient occurring when the Mach number is increased to supersonic values is in good agreement with the predicted increase, and the measured values remain within the boundaries of laminar and turbulent skin friction. The calculated variation of the pressure drag component of the minimum drag coefficient of the wing at supersonic Mach numbers is based upon the results of reference 9, which apply specifically to symmetrical double-wedge sections. Justification for the application of the results of reference 9 to the rounded leading-edge profile of the present model may be found in reference 2. The correspondence of the present results with those of reference 2 at 1.53 Mach number and a somewhat higher Reynolds number is fair.\n\nDrag due to lift.- The component of the drag coefficient that is due to lift is related to the maximum lift-drag ratio in the following manner for a wing of symmetrical section:\n\n$$ \\left(\\frac{L}{D}\\right)_{max} = \\frac{1}{2} \\sqrt{\\frac{1}{C_{D_{min}} \\times \\frac{\\Delta C_D}{\\Delta C_L^2}}} $$\n\nwhere it is convenient to consider the drag due to lift in the form of $\\Delta C_D / \\Delta C_L^2$, termed the \"drag-rise factor.\" Thus, it can be seen that the drag-rise factor can influence the maximum lift-drag ratio to the same degree as the minimum drag coefficient. The variation of the measured drag-rise factor is presented in figure 9 with two calculated variations that describe the limiting values of the drag-rise factor at each Mach number. The lower calculated curve is the variation of the minimum values of the drag due to lift, that is, the condition of complete realization of the theoretically available\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:00.226754+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 40, "total_pages": 92, "image_filename": "19930085870_p40.jpg", "text": "NACA RM No. L9D07\n41\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O Cl\n                {□ Cm\nWedge L.E.     {△ Cl\n                {◇ Cm\n.16\n\n.08\nCL\n0\n-.08\n-.16\n-.24\n\n.01\nCm\n0\n-.01\n\n.06\nCD\n.04\n.02\n0.8\n\n6\n4\nL/D\n2\n0\n\n-6 -4 -2 0 2 4 6 8\nα, deg\n\nElliptical L.E. {O CD\n                {□ L/D\nWedge L.E.     {△ CD\n                {◇ L/D\n\n(A) Wing 11. w=1.280; R = 640,000.\nFigure 5. - Concluded.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:01.031793+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 20, "total_pages": 22, "image_filename": "19930085928_p20.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL 19\n\n[Figure: Graph plotting Total-pressure ratio, $H_3/H_0$ (y-axis, 0 to 1.0) against Mass-flow ratio, $m_1/m_0$ (x-axis, 0 to 1.2). The graph contains three groups of curves labeled $M_0=1.36$, $M_0=1.70$, and $M_0=2.01$. Data points are marked with circles (Model A) and squares (Model B). Dashed lines indicate Unsteady flow. An arrow points to the right with the text \"Normal shock inside ducts at greater mass-flow ratios\". A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 5. -Variation of total-pressure ratio with mass-flow ratio for models without slots.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:01.447888+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 3, "total_pages": 17, "image_filename": "19930085988_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM L9H30\n\nare of a continuous nature from high subsonic to supersonic speeds and\nat high Reynolds numbers.\n\nThis paper presents zero-lift drag data for a body alone and for a\nwing-body configuration having wings of $45^\\circ$ sweepback on the quarter-\nchord line, aspect ratio 4, taper ratio 0.6, and an NACA 65A006 airfoil\nsection in the free-stream direction. The body had a fineness ratio 10\nwith frontal area 6.06 percent of the wing area.\n\nThe Mach number range of the tests was from 0.83 to 1.92 and the\nReynolds number varied from $6 \\times 10^6$ to $23 \\times 10^6$ based on the wing mean\naerodynamic chord.\n\nSYMBOLS\n\n| | |\n| :--- | :--- |\n| $C_D$ | drag coefficient $\\left(\\frac{\\text{Drag}}{qS}\\right)$ |\n| $C_{pb}$ | base-pressure coefficient $\\left(\\frac{p_b - p}{q}\\right)$ |\n| $p_b$ | pressure acting on base of model, pounds per square foot |\n| $p$ | free-stream static pressure, pounds per square foot |\n| $q$ | dynamic pressure, pounds per square foot $\\left(\\frac{1}{2}\\rho V^2\\right)$ |\n| $\\rho$ | air density, slugs/feet$^3$ |\n| $V$ | velocity, feet per second |\n| $M$ | Mach number $\\left(\\frac{V}{c}\\right)$ |\n| $c$ | speed of sound, feet/second |\n| $S_w$ | wing-plan-form area (including area within body), 15.208 square feet |\n| $S_B$ | body frontal area, 0.923 square foot |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:02.547176+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 62, "total_pages": 62, "image_filename": "19930082918_p62.jpg", "text": "NACA TN 1940\n73\n\n[Figure: Micrograph showing a cracked, granular material structure.]\n\n(a) Aged 0.5 hour: failed in 1.81 hours\nunder a stress of 70,000 psi.\n\n[Figure: Micrograph showing a cracked, granular material structure.]\n\n(b) Aged 100 hours: failed in 1.25 hours\nunder a stress of 70,000 psi.\n\nFigure 21.- Effect of aging at $1600^\\circ$ F on fracture characteristics at\n$1200^\\circ$ F of low-carbon N-155 alloy solution-treated 10 hours at\n$2200^\\circ$ F and water-quenched.\n\nNACA-Langley - 8-22-49 - 900", "timestamp": "2026-07-22T06:50:04.059237+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 17, "total_pages": 92, "image_filename": "19930085951_p17.jpg", "text": "NACA RM L9D29\nUNCLASSIFIED\nCONFIDENTIAL\n15\n\nTABLE I\n\nRANGE OF BLADE ANGLE AND ROTATIONAL SPEED\nFOR NACA PROPELLER TESTS\n\n| Figure | Rotational speed (rpm) | Blade angle at 0.75 radius, $\\beta_{0.75R}$ (deg) | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | 20 | 25 | 30 | 35 | 40 | 45 | 50 | 55 |\n| **NACA 10-(3)(062)-045A propeller** | | | | | | | | | |\n| 8 | 1140 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 9 | 1350 | | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 10 | 1500 | | | | | | 45 | | |\n| 11 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 12 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 13 | 2160 | 20 | 25 | 30 | | | | | |\n| 14, 15, 16 | Varied | | | | | | 45 | | |\n| **NACA 10-(3)(05)-045 propeller** | | | | | | | | | |\n| 17 | 1140 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | 55 |\n| 18 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 19 | 1500 | | | | | | 45 | | |\n| 20 | 1600 | 20 | 25 | 30 | | 40 | 45 | | |\n| 21 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 22 | 2160 | 20 | 25 | 30 | | | | | |\n| 23, 24, 25 | Varied | | | | | | 45 | | |\n| **NACA 10-(3)(062)-045 propeller (reference 7)** | | | | | | | | | |\n| 4 | 1140 | | 25 | 30 | 35 | 40 | 45 | 50 | 55 |\n| 5 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 6 | 1500 | | | | | | 45 | | |\n| 7 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 8 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 9 | 2160 | 20 | 25 | 30 | | | | | |\n| 10 | Varied | | | | | | 45 | | |\n| **NACA 10-(3)(08)-045 propeller (reference 7)** | | | | | | | | | |\n| 11 | 1140 | | 25 | 30 | 35 | 40 | 45 | 50 | 55 |\n| 12 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 13 | 1500 | | | | | | 45 | | |\n| 14 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 15 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 16 | 2160 | 20 | 25 | 30 | | | | | |\n| 17 | Varied | | | | | | 45 | | |\n\nUNCLASSIFIED\nCONFIDENTIAL\nNACA", "timestamp": "2026-07-22T06:50:05.787886+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 19, "total_pages": 47, "image_filename": "19930085919_p19.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:50:06.383489+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 27, "total_pages": 52, "image_filename": "19930085911_p27.jpg", "text": "CONFIDENTIAL\n\nFuel regulator\nEvacuated chamber\nFree-stream total pressure\nSynthetic-rubber fuel cell\nHigh-pressure helium\nRegulated helium pressure\nTo 7 fuel nozzles\nTo 12 fuel nozzles\nTo 10 fuel nozzles\nSpiraled helium coils\nFuel-flow orifice\nSpring-loaded valves\nNACA\nFuel tank\n\nFigure 4. - Schematic diagram of fuel system for supersonic 16-inch ram-jet unit.\n\nCONFIDENTIAL\nNACA RM E9F22\n26\n202-172\n2ST2", "timestamp": "2026-07-22T06:50:07.839671+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 3, "total_pages": 132, "image_filename": "19930085990_p3.jpg", "text": "NACA RM A9I01 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nINVESTIGATION OF A THIN WING OF ASPECT RATIO 4 IN THE AMES\n12-FOOT PRESSURE WIND TUNNEL. V - STATIC LONGITUDINAL\nSTABILITY AND CONTROL THROUGHOUT THE SUBSONIC SPEED\nRANGE OF A SEMISPAN MODEL OF A SUPERSONIC AIRPLANE\n\nBy Ben H. Johnson, Jr., and Francis W. Rollins\n\nSUMMARY\n\nWind-tunnel tests have been made of a semispan model of a hypothetical supersonic airplane to determine the static longitudinal-stability and -control characteristics of the airplane throughout the range of subsonic Mach numbers up to 0.95. The semispan model had a long slender fuselage and a wing and horizontal tail of aspect ratio 4 and taper ratio 0.5. The midchord lines of the wing and of the horizontal tail were normal to the plane of symmetry. The profile of the wing and of the tail was a sharp-edged, faired, symmetrical double wedge with a thickness-chord ratio of 0.042. Tests were made with the horizontal tail mounted in the extended wing-chord plane and alternately 69.6 percent of the wing mean aerodynamic chord above the extended wing-chord plane. At a constant Reynolds number of 2,000,000 measurements were made with various stabilizer angles of the lift, drag, and pitching moment of the model at Mach numbers from 0.20 to 0.95. With the wing flaps deflected for maximum lift, similar measurements were made at a Mach number of 0.20 with Reynolds numbers up to 10,000,000. Measurements were made of the dynamic pressure at the two locations of the horizontal tail and of the character and location of the wing wake for the range of Mach numbers and Reynolds numbers noted above.\n\nAt zero lift, the Mach number for drag divergence, defined as the Mach number at which the slope of the drag coefficient with respect to Mach number equals 0.10, was about 0.92 for either location of the horizontal tail. The angle of attack for a constant lift coefficient decreased slightly with increasing Mach number but no marked or abrupt compressibility effects were evident at lift coefficients less than 0.6.\n\nThe contribution of the horizontal tail to the static longitudinal stability at low lift coefficients decreased with increasing Mach number,\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:08.821636+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 6, "total_pages": 46, "image_filename": "19930085983_p6.jpg", "text": "4\nCONFIDENTIAL\nNACA RM A9I27\n\n$\\delta_{Tu}$\narithmetic sum of positive and negative elevon deflections\nuncorrected for angular distortion due to load, degrees\n\n$\\mu$\ncoefficient of viscosity of air, slugs per foot-second\n\n$\\rho$\nmass density of air, slugs per cubic foot\n\nMODEL AND APPARATUS\n\nThe model used in this investigation was the one used in the tests\nreported in reference 6. Photographs of the model are presented in\nfigure 1 and dimensions are given in figures 2 and 3.\n\nThe wing had a leading-edge sweepback of $63^\\circ$, a taper ratio of 0.25,\nand an aspect ratio of 3.5. The streamwise airfoil sections had the\nNACA 64A005 thickness distribution combined with a = 1 mean camber lines.\nThe wing, as developed theoretically by the method given in reference 7,\nwas cambered and twisted to support a uniform distribution of lift over\nits surface at a lift coefficient of 0.25 and a Mach number of 1.5. To\nprovide for twisting of the wing under aerodynamic loads, the model wing\nwas constructed with less twist than was indicated by theory, as is\ndescribed in reference 6.\n\nThe elevons were of constant chord and extended over the outer 50\npercent of the span. Each elevon was supported by three hinges and was\nrestrained near the inner extremity. The ratio of elevon chord to wing\nchord was 1 to 4 at the wing midsemispan. The elevons had radius noses\nwith no aerodynamic balance. The nose gaps were approximately 3/64 inch\nand were unsealed. These large gaps were necessary to permit the desired\nangular deflection since the elevons had considerable spanwise curvature.\nHinge moments were measured by means of a wire-resistance strain gage\nmounted on the restraining member of the elevon on the left-hand wing.\n\nThe model was sting mounted, and the angle of attack was continu-\nously controllable from a remote station during wind-tunnel operation.\nForces and moments acting upon the model were measured by means of a\nwire-resistance strain-gage balance enclosed by the fuselage.\n\nTESTS\n\nLift, drag, pitching-moment, rolling-moment, and elevon-hinge-\nmoment data have been obtained throughout an angle-of-attack range of\n$-3^\\circ$ to $+19^\\circ$. This range was more limited at the larger elevon deflec-\ntions and higher Mach numbers where vibration of either the model or its\nsupport or wind-tunnel power limits were critical. All tests were made\nat an angle of sideslip of $0^\\circ$. The elevons were deflected negatively\nfor longitudinal control and differentially for lateral control as given\nin the following table:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:12.426222+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 55, "total_pages": 72, "image_filename": "19930085491_p55.jpg", "text": "54\n\n$M_o = 1.53$\n\n$\\Lambda_{LE} = 69.9^\\circ$\n\n$R = 0.62 \\text{ million}$\n\nCONFIDENTIAL\n\nPitching-moment coefficient, $C_{m_{\\frac{1}{4}}}$\n\nLift coefficient, $C_L$\n\nAngle of attack, $\\alpha$, deg\n\nDrag coefficient, $C_D$\n\nLift-drag ratio, $\\frac{L}{D}$\n\nLift coefficient, $C_L$\n\nCONFIDENTIAL\n\nNACA RM No. A8J04\n\nWF-70\n\n(e) WF-70\n\nFigure 7- Continued", "timestamp": "2026-07-22T06:50:14.794722+00:00"}
{"citation_id": "19930085975", "source_url": "https://ntrs.nasa.gov/api/citations/19930085975/downloads/19930085975.pdf", "page_number": 8, "total_pages": 30, "image_filename": "19930085975_p8.jpg", "text": "6\nCONFIDENTIAL\nNACA RM L9E10\n\nSubstituting in equation (1) and dividing through by $qSb$ gives\n$$C_{l\\delta_{test}} \\delta = C_{l\\delta} \\delta - C_{l\\delta_{test}} \\delta \\left( C_{l_p} \\frac{\\Delta \\xi}{\\Delta L} qSb \\right) \\quad (4)$$\n\nDividing through by $\\delta$ and transposing terms gives\n$$C_{l\\delta} = C_{l\\delta_{test}} \\left[ 1 + C_{l_p} \\left( \\frac{\\Delta \\xi}{\\Delta L} \\right) qSb \\right] \\quad (5)$$\n\nor\n$$C_{l\\delta} = K C_{l\\delta_{test}} \\quad (6)$$\n\nwhere\n$$K = 1 + C_{l_p} \\left( \\frac{\\Delta \\xi}{\\Delta L} \\right) qSb \\quad (7)$$\n\nEquation (7) involves the use of the $C_{l_p}$ for the rigid wing which has not been determined. However, a first approximation of $C_{l_p}$ can be obtained by\n$$C_{l_{p_{test}}} = - \\frac{C_{l\\delta_{test}}}{\\left( \\frac{pb}{2V} \\right)_{\\delta}}$$\n\nBy using this value of $C_{l_{p_{test}}}$ in equations (7) and (6) a first approximation of $K$ and $C_{l\\delta}$ can be obtained. With this value of $C_{l\\delta}$, a second approximation of $C_{l_p}$ is found and thus, by successive approximations, the final value of $K$ was determined (fig. 5).\n\nReduction of Data\n\nBy use of the correction factor $K$ developed in the previous section, the aileron-effectiveness parameter $C_{l\\delta}$ and the coefficient of damping in roll $C_{l_p}$ were evaluated as follows:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:15.219428+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 87, "total_pages": 98, "image_filename": "19930086073_p87.jpg", "text": "NACA RM A9H04\n85\n\n<!-- Image (108, 110, 885, 997) -->\n\nLift coefficient, $C_L$\nDrag coefficient, $C_D$\n\nAngle of sideslip, $\\beta$, deg\n$\\circ$ 0.0\n$\\square$ 12.0\n\n(b) $C_L$ vs $C_D$.\n\nFigure 19. - Continued.", "timestamp": "2026-07-22T06:50:15.219622+00:00"}
{"citation_id": "19930085991", "source_url": "https://ntrs.nasa.gov/api/citations/19930085991/downloads/19930085991.pdf", "page_number": 1, "total_pages": 24, "image_filename": "19930085991_p1.jpg", "text": "N62 60214\nRM L9I28\n\nNACA RM L9I28\n\nNACA CASE FILE\nCOPY\n\nRESEARCH MEMORANDUM\n\nAN EMPIRICAL CRITERION FOR FIN STABILIZING\nJETTISONABLE NOSE SECTIONS OF AIRPLANES\n\nBy Stanley H. Scher\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\n\nDecember 8, 1949\nDeclassified August 18, 1954", "timestamp": "2026-07-22T06:50:15.220457+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 86, "total_pages": 114, "image_filename": "19930086061_p86.jpg", "text": "82\nNACA RM L9J07\n\n<!-- Image (92, 145, 860, 837) -->\n\nFigure 31.- Effect of $\\alpha$ on the spanwise $c_l$ variation of wing 1; $\\psi = 0^\\circ$.", "timestamp": "2026-07-22T06:50:16.673460+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 55, "total_pages": 60, "image_filename": "19930085862_p55.jpg", "text": "```markdown\nNACA RM No. L9A07\n53\n\n<!-- Image (159, 119, 906, 785) -->\n\n(a) $C_l$ and $C_n$ against $\\alpha$.\n\nFigure 21.- Effects of 0.10c projection step spoilers on characteristics of wing with drooped-nose flaps and trailing-edge split flaps.\n```", "timestamp": "2026-07-22T06:50:18.014864+00:00"}
{"citation_id": "19930085966", "source_url": "https://ntrs.nasa.gov/api/citations/19930085966/downloads/19930085966.pdf", "page_number": 9, "total_pages": 55, "image_filename": "19930085966_p9.jpg", "text": "8\nCONFIDENTIAL\nNACA RM L9B17\n\nscreens on the nozzle exit. The total pressure at station 1 was constant across the section except for a negligible area adjacent to the wall.\n\nThe total-pressure readings from the two rakes at station 6 (see fig. 7) were used to determine a ratio of the weighted average total pressure to the center of the passage total pressure. The average total pressure at the shroud exit was taken to be the product of this ratio and the arithmetic average of the center of the passage tube readings. Average measured pressures and temperatures were plotted against fuel flow and percent of maximum blower speed. All calculations were made using faired values from these curves, examples of which are presented in figures 9 and 10. Figure 9(a) shows average measured diffuser-exit static pressures taken from runs made with constant fuel flow and varying blower speed. Similar data taken at constant blower speed and variable fuel flow are shown in figure 9(b), which also shows solid points taken from faired curves of figure 9(a). The agreement between the curves and the solid points indicates the relative value of faired data obtained by the two methods. Similar data and comparisons are given for nozzle-exit total pressure in figures 10(a) and 10(b).\n\nThe following are relations used in the computations:\n\nVelocity\n$$V = \\sqrt{\\frac{2\\gamma gRT_t}{(\\gamma - 1)} \\left[ 1 - \\left(\\frac{p}{p_t}\\right)^{\\frac{\\gamma-1}{\\gamma}} \\right]}$$\n\nMass flow at stations 1 and 6\n$$\\frac{W}{g} = Ap \\sqrt{\\frac{2\\gamma}{gR(\\gamma - 1)T_t} \\left(\\frac{p_t}{p}\\right)^{\\frac{\\gamma-1}{\\gamma}} \\left[ \\left(\\frac{p_t}{p}\\right)^{\\frac{\\gamma-1}{\\gamma}} - 1 \\right]}$$\n\nDynamic pressure at stations 2 and 3\n$$(p_t - p) = p \\left\\{ \\left[ \\frac{1}{2} + \\sqrt{\\frac{gRT_t(\\gamma - 1)}{2\\gamma p^2} \\left(\\frac{W}{gA}\\right)^2 + \\frac{1}{4}} \\right]^{\\frac{\\gamma}{\\gamma-1}} - 1 \\right\\}$$\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:20.216916+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 41, "total_pages": 92, "image_filename": "19930085870_p41.jpg", "text": "42\nNACA RM No. L9D07\n\nCONFIDENTIAL\n\n.24\nElliptical L.E. {O CL, □ Cm\nWedge L.E. {△ CL, ◇ Cm\n.16\n.08\nCL\n0\n-.08\n-.16\n-.24\n.01\nCm\n0\n-.01\n\n.06\nElliptical L.E. {O CD, □ L/D\nWedge L.E. {△ CD, ◇ L/D\n.04\nCD\n.02\n0.8\n-6\n-4\n-2\n0\n2\n4\n6\n8.0\nα, deg\n6\n4\nL/D\n2\n0\nNACA\n\n(a) Wing 1. w=0.287, R = 1,250,000.\nFigure 6. - Aerodynamic characteristics of 8-percent-thick triangular wings at M=1.92.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:22.745182+00:00"}
{"citation_id": "19930085970", "source_url": "https://ntrs.nasa.gov/api/citations/19930085970/downloads/19930085970.pdf", "page_number": 9, "total_pages": 30, "image_filename": "19930085970_p9.jpg", "text": "NACA RM A9E09 CONFIDENTIAL 7\n\nleading-edge thrust. At subsonic Mach numbers, this curve was obtained from reference 6 and the values are very nearly equal to $1/\\pi A$, the value for a wing with an elliptic span loading. At supersonic Mach numbers above 1.43, the calculated optimum drag-rise factor was obtained by the methods of reference 7. The variation of the values between Mach numbers of 1.43 and 1.0 has been represented by a straight line. The use of $1/\\pi A$ as the value of the drag-rise factor at a Mach number of unity can be justified by the analysis of reference 1. The upper curve represents the drag-rise factor for the case of zero leading-edge thrust at each Mach number. Since this case corresponds to the condition for which the resultant force acts normal to the chord line, the curve has been determined as the variation of the reciprocal of the experimental lift-curve slope. At subsonic Mach numbers, it is seen from the results shown in figure 9 that the available leading-edge thrust was not completely realized on the model. It is believed that this loss of leading-edge thrust was caused by flow separation near the leading edges that occurred at the low test Reynolds numbers. It is also noted that the measured drag-rise-factor variation virtually parallels the upper curve up to Mach numbers of about 1.2, but at the higher supersonic Mach numbers the experimental results approach the calculated lower limiting curve.\n\nThe cause of the discrepancy between the result of this report at the highest Mach number and that of reference 2 is not known.\n\nMaximum Lift-Drag Ratio\n\nIn figure 10, the measured variation of the maximum lift-drag ratio with Mach number is compared with the experimental results of references 2 and 3 and the calculated variations. The latter variations are based upon the calculated minimum drag coefficients shown in figure 8 and the calculated optimum drag-rise factors shown in figure 9. By reference to figure 8, where the measured minimum drag coefficients are seen to fall well within the calculated limits, it is deduced that a major portion of the difference between the present measured and calculated maximum lift-drag ratios is due to the high values of drag due to lift observed at the low test Reynolds numbers. Furthermore, about 60 percent of the difference between the present maximum lift-drag ratios and those of reference 3 can be traced to the improved drag-rise factors accompanying the higher Reynolds number of the latter investigation. The closer agreement of the present results with calculated values at the higher supersonic Mach numbers is primarily a reflection of the corresponding trend of agreement shown by the measured drag-rise factor. The lack of agreement between the present result and that of reference 2 largely results\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:24.869870+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 1, "total_pages": 32, "image_filename": "19930085992_p1.jpg", "text": "NACA RM L9E17\n\nNACA\n\nRESEARCH MEMORANDUM\n\nPRELIMINARY EXPERIMENTAL INVESTIGATION OF EFFECTS OF\nAERODYNAMIC SHAPE OF CONCENTRATED WEIGHTS ON\nFLUTTER OF A STRAIGHT CANTILEVER WING\n\nBy\n\nJohn L. Sewall and Donald S. Woolston\n\nLangley Aeronautical Laboratory\nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\n\nWASHINGTON\nJuly 18, 1949", "timestamp": "2026-07-22T06:50:26.044600+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 28, "total_pages": 52, "image_filename": "19930085911_p28.jpg", "text": "NACA RM E9F22\nCONFIDENTIAL\n27\n\n[Figure: (a) Three-quarter front view. NACA C-20036 11-20-47]\n\n[Figure: (b) Three-quarter rear view. NACA C-20035 11-20-47]\n\nFigure 5. - Flame holder for supersonic ram-jet units 16-A-2, 16-A-3, and 16-A-4.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:26.248816+00:00"}
{"citation_id": "19930085979", "source_url": "https://ntrs.nasa.gov/api/citations/19930085979/downloads/19930085979.pdf", "page_number": 5, "total_pages": 25, "image_filename": "19930085979_p5.jpg", "text": "4\nNACA RM E9E12\n\nfunction of tunnel-air velocity, angle of attack, gas flow, and\ngas temperature. The investigation was conducted at a fixed tail-\ncone position corresponding to an inlet-velocity ratio of 0.89\nwithout bleedback and at a free-stream total temperature of 0° F\nfor tunnel-air velocities from 200 to 450 feet per second and\nangles of attack from 0° to 8°. Gas flows and plenum-chamber-gas\ntemperatures ranged from 0.59 to 1.55 pounds per second and from\n600° to 1000° F, respectively. For each plenum-chamber-gas\ntemperature, the gas pressure was varied from 3000 to 6000 pounds\nper square foot absolute.\n\nIcing with hot-gas bleedback. - An investigation to determine\nthe critical icing criterion as a function of mass flow, gas flow,\nangle of attack, and liquid-water content for a constant free-\nstream total temperature of 0° F was conducted in a manner similar\nto that of reference 1. This investigation was conducted at tunnel-\nair velocities of 200, 275, 355, and 435 feet per second at an\nangle of attack of 0°. The liquid-water content ranged from 0.70 to\n1.4 grams per cubic meter at an average drop diameter of 15 microns.\nThe range of gas flows and plenum-chamber-gas temperatures was the\nsame as those employed for the aerodynamic investigation with hot-\ngas bleedback.\n\nRESULTS AND DISCUSSION\n\nAerodynamic Investigation without Bleedback\n\nMass-flow characteristics. - The mass flow through the model\nincreased nearly linearly with tunnel-air velocity for a fixed tail-\ncone position and angle of attack. A maximum flow of approximately\n32.1 pounds per second was obtained at an inlet-velocity ratio of\n0.89, a tunnel-air velocity of 460 feet per second, and an angle\nof attack of 0°.\n\nRam-pressure recovery. - A ram-pressure recovery of approxi-\nmately 0.99 was obtained at an angle of attack of 0° and an inlet-\nvelocity ratio of 0.89. Ram-pressure recovery $\\eta$ is defined as\n\n$$ \\eta = 1 - \\frac{P_0 - P_F}{q_0} $$\n\nwhere\n\n$P_0$ free-stream total pressure, pounds per square foot absolute", "timestamp": "2026-07-22T06:50:26.605501+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 21, "total_pages": 22, "image_filename": "19930085928_p21.jpg", "text": "Normal shock wave\nNormal shock wave\n\nCONFIDENTIAL\n\nMaximum total-pressure ratio, $(H_3/H_0)_{max}$\n\nMaximum total-pressure ratio, $(H_3/H_0)_{max}$\n\nModel of reference 1\n$\\circ$ .085in.X.300in. slots\nModel A\n$\\square$ .085in.X.300in. slots\n$\\diamond$ .085in.X.450in. slots\n\nModel of reference 1\n$\\circ$ .085in.X.300in. slots\nModel B\n$\\square$ .085in.X.300in. slots\n$\\diamond$ .085in.X.450in. slots\n$\\triangle$ .044in.X.450in. slots\n\nNACA\n\nMach number, $M_0$\n\nMach number, $M_0$\n\nFigure 6.—Variation of maximum total-pressure ratio with Mach number for models with slots.\n\nCONFIDENTIAL\n\nNACA RM No. A9A31\n\n20", "timestamp": "2026-07-22T06:50:26.804562+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 20, "total_pages": 47, "image_filename": "19930085919_p20.jpg", "text": "```markdown\nNACA RM No. A9C21\n\nCONFIDENTIAL\n\nHinge line of 25-percent-\nchord elevon\n\nConstant along span\n\n3.43\n0.082\n\n0.25c\n\nStation 30.00\n\n0.06\n\n8.57\n\n0.75c\n\nDimensions in inches\n\nLong fuselage\nnose\n\nShort\nfuselage nose\n\n63°\n\n20.40\n\n36.17\n\n54.86\n\n1.85\n\n204.00\n\nSection A-A\n\narea of semispan wing 14.274 sq ft\narea of constant-chord elevon 1.782 sq ft\narea of 0.25-chord elevon 1.248 sq ft\n\n0.38c\n\n1.372\n\nHinge line of\nconstant-chord\nelevon\n\n5.28\nconstant\n\n0.75c\n\n8.57\n\nM.A.C.\n38.41\n\n24.00\n\n0.06\n\n30.00\n\nShort fuselage\ntail\n\nLong\nfuselage\ntail\n\n13.20\n\nMoment center, 0.25c\n\n60.00\n\nNACA\n\nFigure 2.- The model geometry.\n\nCONFIDENTIAL\n\n19\n```", "timestamp": "2026-07-22T06:50:30.350585+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 56, "total_pages": 72, "image_filename": "19930085491_p56.jpg", "text": "$M_o = 1.53$\n$R = 0.62 \\text{ million}$\n(characteristics based on\nwing dimensions of WF-63)\n\nCONFIDENTIAL\n\nLift coefficient, $C_L$\nAngle of attack, $\\alpha$, deg\n\nPitching-moment coefficient, $C_{m_{\\frac{1}{4}}}$\n\nDrag coefficient, $C_D$\nAngle of attack, $\\alpha$, deg\n\n(f) Fuselage alone\n\nNACA\n\nFuselage alone\n\nCONFIDENTIAL\n\nNACA RM No. A8J04\n\nFigure 7- Concluded.\n\n55", "timestamp": "2026-07-22T06:50:32.816959+00:00"}
{"citation_id": "19930085988", "source_url": "https://ntrs.nasa.gov/api/citations/19930085988/downloads/19930085988.pdf", "page_number": 4, "total_pages": 17, "image_filename": "19930085988_p4.jpg", "text": "```markdown\nNACA RM L9H30 CONFIDENTIAL 3\n\nMODELS AND TESTS\n\nThe general arrangements and profile coordinates for the test configuration are shown in figure 1 and table I, and photographs of the test models on the launching stand are given as figure 2. The body was identical for both configurations and had a length of 10.8 feet, diameter of 1.08 feet, and frontal area of 0.923 square foot. The body shape was modified from that of the free-fall bodies, reference 1, by cutting off the pointed stern at the 83.5-percent station. A base-pressure tube was located in the stern end of the body; a detail of its installation is shown in figure 3. The wing had a sweepback of $45^\\circ$ on the quarter-chord line, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil sections parallel to the model center line. The wing-plan-form area was 15.208 square feet and the wing was located such that the one-quarter point of the mean aerodynamic chord fell at the station corresponding to the maximum diameter of the body (6.5 feet rearward of the nose). The wingless configuration was stabilized by four fins and the winged configuration by two fins in the vertical plane and by the wing in the horizontal plane. All fins were of 1.23 square feet exposed area each, having approximately $60^\\circ$ sweepback and mean thickness ratio of 3 percent.\n\nWith the exception of the metal fins, all surfaces of both configurations were wood and had a smooth and highly polished lacquered finish.\n\nThe wingless and winged configurations were each propelled by a Deecon rocket motor which delivered approximately 6200 pounds of thrust for 3.2 seconds.\n\nVelocity and drag were obtained from the CW Doppler velocimeter described in reference 2 and drag and base pressure were reduced from data telemetered by a two-channel instrumentation unit incorporating a longitudinal accelerometer and pressure cell. Trajectory and atmospheric data were obtained from the NACA modified SCR-584 radar tracking unit and by radiosonde observations.\n\nTotal-drag coefficients refer to the measured total drag of the test configurations and base-drag coefficients refer to the drag contribution of the base. The base-drag coefficient is computed as equal to the product of the base-pressure coefficient and the ratio of the base area to wing area (0.015) by assuming that the measured base pressure acts over the entire area of the base.\n\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:50:33.830856+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 87, "total_pages": 114, "image_filename": "19930086061_p87.jpg", "text": "```markdown\nNACA RM L9J07\n83\n\n<!-- Image (139, 141, 901, 829) -->\n\nFigure 32.- Effect of $\\alpha$ on the spanwise $c_l$ variation of wing 2; $\\psi = 0^\\circ$.\n```", "timestamp": "2026-07-22T06:50:37.005847+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 18, "total_pages": 92, "image_filename": "19930085951_p18.jpg", "text": "```markdown\n16\nUNCLASSIFIED\nCONFIDENTIAL\nNACA RM L9D29\n\nTABLE I\nRANGE OF BLADE ANGLE AND ROTATIONAL SPEED\nFOR NACA PROPELLER TESTS - Concluded\n\n| Figure | Rotational speed (rpm) | Blade angle at 0.75 radius, $\\beta_{0.75R}$ (deg) | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | | | | | | | |\n| **NACA 10-(3)(08)-03 propeller (reference 3)** | | | | | | | | | |\n| 19 | 1140 | | | 30 | 35 | 40 | 45 | 50 | 55 |\n| 20 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 21 | 1500 | | | | | | 45 | | |\n| 22 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 23 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 24 | 2160 | 20 | 25 | 30 | | | | | |\n| 25 | Varied | | | | | | 45 | | |\n| | | | | | | | | | |\n| **NACA 10-(3)(08)-03R propeller (reference 5)** | | | | | | | | | |\n| 3 | 1140 | | | | 35 | 40 | 45 | 50 | 55 |\n| 4 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 8 | 1500 | | | | | | 45 | | |\n| 5 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 6 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 7 | 2160 | 20 | 25 | 30 | | | | | |\n| 9, 10 | Varied | | | | | | 45 | 50 | |\n| | | | | | | | | | |\n| **NACA 10-(3)(12)-03 propeller (reference 8)** | | | | | | | | | |\n| 2 | 1140 | | 25 | 30 | 35 | 40 | 45 | 50 | 55 |\n| 3 | 1350 | 20 | 25 | 30 | 35 | 40 | 45 | 50 | |\n| 4 | 1500 | | | | | | 45 | | |\n| 5 | 1600 | 20 | 25 | 30 | 35 | 40 | 45 | | |\n| 6 | 2000 | 20 | 25 | 30 | 35 | | | | |\n| 7 | 2160 | 20 | 25 | 30 | | | | | |\n| 8 | Varied | | | | | | 45 | | |\n\nNACA\nUNCLASSIFIED\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:50:37.207991+00:00"}
{"citation_id": "19930085983", "source_url": "https://ntrs.nasa.gov/api/citations/19930085983/downloads/19930085983.pdf", "page_number": 7, "total_pages": 46, "image_filename": "19930085983_p7.jpg", "text": "NACA RM A9I27 CONFIDENTIAL 5\n\n| Elevon deflection angles | | | |\n| :--- | :--- | :--- | :--- |\n| **Longitudinal-control data** | | **Lateral-control data** | |\n| $\\delta_{L_u}$ (deg) | $\\delta_{R_u}$ (deg) | $\\delta_{L_u}$ (deg) | $\\delta_{R_u}$ (deg) |\n| 0 | 0 | 0 | 0 |\n| -5 | -5 | 10 | -10 |\n| -10 | -10 | 20 | -20 |\n| -15 | -15 | 30 | -30 |\n| -20 | -20 | - | - |\n| -25 | -25 | - | - |\n\nThe tests were performed at several Mach numbers ranging from 0.20 to 0.93 at a constant Reynolds number of 2.0 million.\n\nCORRECTIONS\n\nThe data have been corrected for the effects of tunnel-wall interference, constriction due to the tunnel walls, base pressure, and static tares due to the weight of the model. No correction has been applied to account for the change of elevon deflection under load upon the force and moment coefficients except when presented as functions of elevon angle. The angle of attack of the model was measured visually by means of a cathetometer; hence, no corrections were necessary to account for deflection of the support equipment. Precision of the force and moment measurements obtained from the strain-gage balance has been discussed in reference 6.\n\nTunnel-Wall Interference\n\nCorrections to the data to account for induced tunnel-wall interference have been determined by the method of Glauert (reference 8). Since the ratio of model span to tunnel diameter was small, the total corrections were small, and no account was taken of sweepback or of the differential flap deflections. The following corrections were added:\n\n$$ \\Delta\\alpha = 0.26 \\ C_L $$\n$$ \\Delta C_D = 0.0046 \\ C_L^2 $$\n\nNo correction was applied to the pitching moment.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:38.893068+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 56, "total_pages": 60, "image_filename": "19930085862_p56.jpg", "text": "```markdown\n54\nNACA RM No. L9A07\n\n<!-- Image (188, 109, 810, 904) -->\n\n(b) $C_L$, $C_D$, and $C_m$ against $\\alpha$.\nFigure 21.- Concluded.\n```", "timestamp": "2026-07-22T06:50:41.616556+00:00"}
{"citation_id": "19930085992", "source_url": "https://ntrs.nasa.gov/api/citations/19930085992/downloads/19930085992.pdf", "page_number": 2, "total_pages": 32, "image_filename": "19930085992_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:50:41.811998+00:00"}
{"citation_id": "19930085990", "source_url": "https://ntrs.nasa.gov/api/citations/19930085990/downloads/19930085990.pdf", "page_number": 4, "total_pages": 132, "image_filename": "19930085990_p4.jpg", "text": "2\nCONFIDENTIAL\nNACA RM A9I01\n\nprimarily due to an increase with Mach number of the rate of change of\neffective downwash angle with angle of attack. For the model with the\nhorizontal tail in the extended wing-chord plane, this decrease in the\ncontribution of the horizontal tail to the static longitudinal stability\nwas aggravated by the reduction with increasing Mach number in the\ndynamic-pressure ratio at the tail. With the horizontal tail mounted in\nthe extended wing-chord plane, static longitudinal stability existed\nabout the quarter point of the wing mean aerodynamic chord at all lift\ncoefficients for Mach numbers less than 0.87. At Mach numbers between\n0.87 and 0.95, the model was neutrally stable or unstable at lift\ncoefficients less than 0.30. With the horizontal tail mounted above the\nextended wing-chord plane, the results indicated static longitudinal\nstability at all lift coefficients for all Mach numbers for which data\nwere obtained. For both positions of the tail, either an all-movable\nstabilizer or a constant-chord elevator provided sufficient longitudinal\ncontrol to balance the airplane at all test Mach numbers.\n\nINTRODUCTION\n\nAs a part of a general program to determine the subsonic character-\nistics of wing plan forms suitable for flight at supersonic speeds, a\nseries of tests of a thin sharp-edged wing having an aspect ratio of 4\nand a taper ratio of 0.5 have been conducted. The midchord line of the\nwing was normal to the air stream. Results of these tests have been\nreported in references 1 through 4. Results of tests at transonic speeds\nof a wing of identical plan form and similar profile have been reported in\nreference 5.\n\nThe purpose of the present report is to summarize the wing data in\nterms of the static longitudinal-stability and -control characteristics\nthroughout the subsonic speed range of a hypothetical airplane employing\nthis wing. The airplane was represented by a semispan model comprising\nthe wing, a slender pointed fuselage, and a horizontal tail geometrically\nsimilar to the wing. Force and moment characteristics of the wing, of the\nwing-fuselage combination, and of the complete model with two different\ntail heights are presented for Mach numbers up to 0.95 and a Reynolds\nnumber of 2,000,000. With the flaps on the wing deflected for maximum\nlift, similar data are presented for a Mach number of 0.20 and Reynolds\nnumbers up to 10,000,000. The dynamic pressure at the horizontal tail\nand the location of the wing wake are presented for the wing-fuselage\ncombination for the same ranges of Reynolds number and Mach number. The\ntests of the wing-tail-fuselage combinations were conducted with various\nhorizontal-stabilizer settings to investigate the longitudinal control\nafforded by an all-movable horizontal tail. Data for an identical hori-\nzontal tail with a constant-chord elevator (reference 6) have been used\nwith the wing-fuselage data to calculate the longitudinal-control\ncharacteristics of the model with a fixed stabilizer and an elevator.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:50:44.731785+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 51, "total_pages": 149, "image_filename": "19930083192_p51.jpg", "text": "NACA TN 1976\n47\n\nMethods\n\nThe analytical method developed in reference 38 was to reduce the airplane structure to an \"equivalent\" biplane whose upper wing had the same motion as the original wing tip and, by using an effective damping factor instead of unsteady-lift functions, to obtain two simultaneous linear differential equations. The equivalent biplane, in which the upper wing is connected by springs to the lower wing-fuselage combination, is adjusted so that the components have the same motions as the wing tip and fuselage of the airplane under study. The equivalent system must include the proper distribution of aerodynamic forces as well as inertia and elastic forces. The deflection of the upper wing of the biplane is taken as a measure of dynamic stress. The dynamic-stress ratio is defined as the ratio of dynamic to static wing-tip deflections. The static deflection is computed in the same manner as for normal design by including inertia effects but neglecting aerodynamic damping due to vibration of the upper wing.\n\nThe spring constant, equivalent masses, and aerodynamic damping for the equivalent system are calculated so that the static and vibration characteristics of the original wing are represented. In general, the following conditions are to be satisfied:\n\n1. The total mass of and the total load on the equivalent biplane should be identical with those of the original airplane.\n\n2. The upper wing should deflect under the equivalent static load the same amount as the original wing tip under its corresponding static-load distribution.\n\n3. The natural frequency of the equivalent system should be the same as that of the original wing.\n\n4. The kinetic energy of vibration of the upper wing should closely approximate that of the original wing for the same tip amplitude of vibration.\n\n5. The damping coefficient of the upper wing should represent, at least up to peak load, the damping of the motion of the original wing.\n\nThe shape of the forcing function used in the calculations of reference 38 was obtained from accelerometer records of gust-tunnel tests of different models by subtracting the computed acceleration increment due to vertical motion. This procedure was followed since the prediction of airplane reactions as affected by stability and other factors is of doubtful accuracy, as previously noted. For representative gust sizes ($H = 0$ to $H = 20$ chords) it was found in reference 38 that a curve of the form $Ate^{-bt}$ was a good representation. The forcing function for", "timestamp": "2026-07-22T06:50:44.953850+00:00"}