PDF/ocr_dataset_1071.jsonl

{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 20, "total_pages": 33, "image_filename": "19930085900_p20.jpg", "text": "NACA RM L9D20\n19\n\nCONFIDENTIAL\n\nStation 10 18 26 34 42.2\n258 jets\n32\n194 jets\n24\n130 jets\n16\n66 jets\n8\n\nSide views of port chine jets\n\nTrim, deg\n10\n8\n6\n4\n2\n0\nO 32-inch chines (258 jets)\n$\\Delta$ 24-inch chines (194 jets)\n$\\square$ 16-inch chines (130 jets)\n$\\diamond$ 8-inch chines (66 jets)\n\nEffective hydrodynamic lift and load on water, lb\n8\n7\n6\n5\n4\n3\n2\n1\n0\n-1\n-2\nLoad on water\n\nResistance, lb\n7\n6\n5\n4\n3\n2\n1\n0\n0 4 8 12 16 20 24 28 32 36 40 44 48 52 56 60\nSpeed, fps\nNACA\n\nFigure 8.- Effect of length of jet rows; $\\frac{1}{4}$-inch-spaced jets normal to center line.\nCONFIDENTIAL", "timestamp": "2026-07-22T06:36:23.618226+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 18, "total_pages": 29, "image_filename": "19930085899_p18.jpg", "text": "```markdown\nNACA RM No. L9J21\n\nM\n1.18 $\\Delta$\n1.15 $\\triangleleft$\n1.10 $\\nabla$\n1.08 $\\diamond$\n1.05 $\\triangledown$\n1.03 $\\diamond$\n1.00 $\\square$\n.98 $\\Delta$\n.95 $\\triangleleft$\n.93 $\\nabla$\n.90 $\\triangledown$\n.88 $\\diamond$\n.85 $\\nabla$\n.80 $\\diamond$\n.70 $\\square$\n.60 $\\circ$\n\nAngle of attack, $\\alpha$, deg\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n0\n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"timestamp": "2026-07-22T06:36:28.824046+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 56, "total_pages": 65, "image_filename": "19930082546_p56.jpg", "text": "NACA TN No. 1870\n55\n\n$M_T$\n\n0.45\n0.60\n0.75\n0.90\n1.00\n\n$\\frac{X}{D} = +\\frac{1}{8}$\n$\\frac{X}{D} = 0$\n[Figure: NACA logo]\n$\\frac{X}{D} = -\\frac{1}{8}$\n\nFigure 18.— Effect of tip Mach number on the pressure wave forms at three different points in space for NACA 4-(5)(08)-03 propeller. B = 2; $\\beta_{0.75} = 12^\\circ$; $\\frac{d}{D} = 0.167$. (Bottom trace in each photograph is 300 cps timing line.)", "timestamp": "2026-07-22T06:36:29.595277+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 5, "total_pages": 47, "image_filename": "19930085919_p5.jpg", "text": "4\nCONFIDENTIAL\nNACA RM No. A9C21\n\nf split flap\ni induced\nn extended-nose flap\nu uncorrected\n\nCORRECTIONS\n\nAn explanation of the method used in calculating the wind-\ntunnel-wall corrections which were applied to the data is given in\nthe appendix. The equations used in correcting the data are as\nfollows:\n\n$$C_D = C_{D_u} + 0.0319 \\quad C_{L_u}^2$$\n\n$$C_L = 0.99 \\quad C_{L_u}$$\n\n$$C_l = 0.938 \\quad C_{l_u}$$\n\n$$C_m = C_{m_u} + 0.0010 \\quad C_{L_u}$$\n\n$$\\alpha = \\alpha_u + 1.36 \\quad C_{L_u} + 0.19 \\quad (C_{L_u})_{\\delta=0^\\circ}$$\n\nMeasurements were made of the deflection due to the aerodynamic\nloads of the model at various spanwise positions and of the change\nin angle of attack of the wing tip for dynamic pressures ranging\nfrom 20 to 150 pounds per square foot. For a lift coefficient of\n0.35 and a dynamic pressure of 50 pounds per square foot the wing\ntip deflected 0.33 inch above its no-load position; however, only a\nnegligible change in angle of attack of the wing tip was measured.\nEvidence that the effects of model distortion were negligible was\nalso obtained from tests of this model in the Ames 12-foot pressure\nwind tunnel (reference 3) at dynamic pressures of 53 and 105 pounds\nper square foot for a constant Reynolds number of $9.75 \\times 10^6$. Only\nsmall effects on the aerodynamic characteristics of the wing were\nproduced by this dynamic-pressure variation. Hence, no corrections\nhave been applied to the data of the present tests for the effects\nof model distortion.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:36:33.349074+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 34, "total_pages": 60, "image_filename": "19930085862_p34.jpg", "text": "32\nNACA RM No. L9A07\n\n$C_L$\n1.4\n1.2\n1.0\n.8\n.6\n.4\n.2\n0\n\n$\\delta_a$\n(deg)\n$\\circ$ -25\n$\\square$ 0\n$\\triangle$ 25\n\n$C_m$\n.04\n0\n-.04\n-.08\n-.12\n-.16\n\n-4 0 4 8 12 16 20 24\n$\\alpha$, deg\n\n(c) $C_L$ and $C_m$ against $\\alpha$.\nFigure 8.- Concluded.", "timestamp": "2026-07-22T06:36:35.610550+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 59, "total_pages": 78, "image_filename": "19930082483_p59.jpg", "text": "NACA TN No. 1807\n57\n\n1032\n\n<!-- Image (148, 199, 899, 808) -->\n\nFigure 3. - Nozzle and rotor-blade detail at pitch line.", "timestamp": "2026-07-22T06:36:35.802585+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 1, "total_pages": 22, "image_filename": "19930085928_p1.jpg", "text": "Copy No. 258\nRM No. A9A31\n\nNACA RM No. A9A31\n\nCONFIDENTIAL\n\nNACA\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION AT SUPERSONIC SPEEDS OF\nTWIN-SCOOP DUCT INLETS OF EQUAL AREA. IV - SOME\nEFFECTS OF INTERNAL DUCT SHAPE UPON AN\nINLET ENCLOSING 37.2 PERCENT OF THE\nFOREBODY CIRCUMFERENCE\n\nBy Wallace F. Davis, Sherman S. Edwards,\nand George B. Brajnikoff\n\nAmes Aeronautical Laboratory\nMoffett Field, Calif.\n\nCLASSIFICATION CHANGED TO UNCLASSIFIED\nAUTHORITY: NACA RESEARCH ABSTRACT NO. 98\nDATE: MARCH 26, 1956\nWHL\n\nThis document contains 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 known loyalty and discretion who of necessity must be informed thereof.\n\nNATIONAL ADVISORY COMMITTEE\nFOR AERONAUTICS\nWASHINGTON\nMarch 15, 1949\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:36:35.862495+00:00"}
{"citation_id": "19930085548", "source_url": "https://ntrs.nasa.gov/api/citations/19930085548/downloads/19930085548.pdf", "page_number": 39, "total_pages": 46, "image_filename": "19930085548_p39.jpg", "text": "38\nNACA RM No. E8L30\n\nCompression\nratio\n2.4\nO 4.00\n4.50\n5.25\n7.00\nPressure\nratio\n2.2\nCalculated\nExperimental\nMaximum combustion pressure\nCompression pressure\n2.0\n1.8\n1.6\n1.4\n1.2\n1.0\n0\n.01\n.02\n.03\n.04\n.05\nFuel-air ratio\nNACA\n\nFigure 16. - Effect of compression ratio and fuel-air\nratio on ratio of maximum combustion pressure to\ncompression pressure. Inlet-manifold temperature,\n400° F; inlet-manifold pressure, 100 pounds per\nsquare inch absolute.", "timestamp": "2026-07-22T06:36:35.862700+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 72, "total_pages": 98, "image_filename": "19930086073_p72.jpg", "text": "70\nNACA RM A9H04\n\nLift coefficient, $C_L$\nPitching-moment coefficient, $C_m$\n\n| Symbol | Right aileron deflection, $\\delta_{a_R}$, deg |\n| :---: | :---: |\n| $\\square$ | +10.8 |\n| $\\circ$ | 0 |\n| $\\diamond$ | -10.8 |\n\n(c) $C_L$ vs $C_m$.\n\nFigure 15.— Continued.", "timestamp": "2026-07-22T06:36:44.572079+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 42, "total_pages": 62, "image_filename": "19930082918_p42.jpg", "text": "NACA TN 1940\n\n45\n\n[Figure: Micrograph showing grain structure with small precipitates]\n\n(e) Aged 10 hours.\n\n[Figure: Micrograph showing grain structure with slightly coarser precipitates and some boundary changes]\n\n(f) Aged 30 hours.\n\n[Figure: Micrograph showing further coarsening of precipitates and more defined grain boundaries]\n\n(g) Aged 100 hours.\n\n[Figure: Micrograph showing significant coarsening, large precipitates, and pronounced grain boundary features]\n\n(h) Aged 1000 hours.\n\nFigure 5.— Concluded.\n\n[NACA logo]", "timestamp": "2026-07-22T06:36:51.048826+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 8, "total_pages": 34, "image_filename": "19930085913_p8.jpg", "text": "NACA RM L9F24\n7\n\n| Model | Phase-angle relationship | | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | **Gage 1** | **Gage 2** | **Gage 3** | **Gage 4** |\n| A | Reference | $180^\\circ$ | $0^\\circ$ | $180^\\circ$ |\n| B | Reference | $180^\\circ$ | $0^\\circ$ | $180^\\circ$ |\n| C | Reference | $0^\\circ$ | $0^\\circ$ | $0^\\circ$ |\n\nThe way in which this table was used is illustrated with the aid of the following sample oscillograph record:\n\n[Figure: Oscillograph record showing four wave traces labeled Gage 1 (Root torsion), Gage 2 (Root bending), Gage 3 (Tip torsion), and Gage 4 (Tip bending). A calibration frequency trace (100 cps) is on the far right. Timing lines indicating 0.10 sec are marked on the first trace.]\n\nIf the record is assumed to be obtained from any one of the models, the phase-angle relationship between the bending and torsional stresses of the model at the strain gages would be obtained as follows:\n\n| Model | Gage 1 | Gage 2 | Gage 3 | Gage 4 |\n| :--- | :--- | :--- | :--- | :--- |\n| A | Reference | $0^\\circ$ | $180^\\circ$ | $0^\\circ$ |\n| C | Reference | $180^\\circ$ | $180^\\circ$ | $180^\\circ$ |\n\nThe gage traces on the sample oscillograph record are numbered and labeled and for all the records of figure 1 and data of table I the identification is the same.", "timestamp": "2026-07-22T06:36:51.269259+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 6, "total_pages": 22, "image_filename": "19930085922_p6.jpg", "text": "NACA RM No. L9C23\n\nThe variation of the damping-in-roll coefficient $C_{l_p}$ with Mach number is presented in figure 11 and the effect of fins is presented in figure 12. It will be noted that the experimental variation of $C_{l_p}$ with Mach number (fig. 11) agrees fairly well with theory (reference 3), although the absolute values of the damping coefficients are slightly greater than the theoretical values. The increase in the magnitude of $C_{l_p}$ with angle of attack, which is evident from figure 11, (particularly at the high Mach numbers) is not accounted for in the basic theory as ordinarily applied. However, experimental data on a similar wing (reference 4) gave a similar variation of $C_{l_p}$ with angle of attack. The slight reduction in magnitude of $C_{l_p}$ at $1.8^\\circ$ angle of attack may be caused by a local change in the section lift-curve slope. Reference 5 indicates that $C_{l_p}$ of a wing is approximately proportional to the slope of the section lift-curve slope at any particular angle of attack. However, since the actual section lift characteristics for this airfoil are not available, no analysis has been made to check this effect.\n\nThe effect of the fins on the damping coefficient was almost negligible (fig. 12). However, the addition of the vertical fins produced a marked increase in the effectiveness of the ailerons in producing both rolling moment and rolling velocity (fig. 10). This effect can probably be attributed to the end-plate effect of the fins.\n\nCONCLUSIONS\n\nThe results of an investigation of the effects of Mach number and angle of attack on the damping-in-roll characteristics of a $35^\\circ$ swept-back wing of aspect ratio 3 and taper ratio 0.6 indicate the following conclusions:\n\n1. The damping-in-roll derivative $C_{l_p}$ increased in magnitude gradually with Mach number in a manner similar to that predicted from theory.\n\n2. The damping-in-roll derivative $C_{l_p}$ generally increased in magnitude with angle of attack within the test range, especially at higher Mach numbers.", "timestamp": "2026-07-22T06:36:51.895271+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 9, "total_pages": 52, "image_filename": "19930085911_p9.jpg", "text": "8 CONFIDENTIAL NACA RM E9F22\n\nThe equations used in determining the free-stream conditions and the performance are given in the appendix.\n\nRESULTS AND DISCUSSION\n\nTime histories of the ram-jet-unit performance are presented in figures 7 to 10. In general, these figures include resultant flight conditions, independent test variables, diffuser conditions, combustion-chamber-inlet variables, and performance variables. The solid curves are measured values and the broken lines represent approximated values. Such approximations were made wherever the telemetering records were so vibratory that no exact data could be measured, although the trend in the data could be clearly established.\n\nUnit A-2 was released at a free-stream Mach number of 0.38 and an altitude of 30,000 feet, as compared with initial free-stream Mach numbers of 0.55 to 0.59 and altitudes of 32,000 to 36,000 feet for the other units. The variation in launching conditions, however, is believed to have had little effect on the final performance of the ram-jet units.\n\nCombustion Performance\n\nIn the comparison of the combustion performance of the four ram-jet units, it should be recalled that the units were similar in construction except for the flame holders.\n\nThe effects of fuel flow and flame-holder design on the free-stream Mach numbers are shown in figure 11. For comparative purposes, fuel flow is shown as a function of free-stream total pressure because this pressure actuated the fuel regulator. The calculated fuel flow necessary to obtain a fuel-air mixture of 0.067 is included. No fuel flow data were obtained for unit A-2.\n\nCombustion occurred in unit A-2 only momentarily. As a result, the maximum free-stream Mach number was only 0.92 shortly before impact. Units A-3 and A-4 had relatively high fuel flows and reached maximum free-stream Mach numbers of 1.34 and 1.25, respectively. Combustion was sporadic throughout the flights with regions of extreme vibration and very low combustion efficiency. Unit A-5 operated at low rates of fuel flow and continuously good combustion with rapidly increasing flight Mach numbers occurred throughout most of the flight. A maximum free-stream Mach number of 1.73 occurred at an altitude of 4900 feet. No data were obtained below\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:36:52.385685+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 21, "total_pages": 33, "image_filename": "19930085900_p21.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:36:54.195593+00:00"}
{"citation_id": "19930085870", "source_url": "https://ntrs.nasa.gov/api/citations/19930085870/downloads/19930085870.pdf", "page_number": 23, "total_pages": 92, "image_filename": "19930085870_p23.jpg", "text": "22 CONFIDENTIAL NACA RM No. L9D07\n\nTABLE 1.- DIMENSIONS OF TRIANGULAR-WING MODELS\n\n(a) 8-Percent-Thick Triangular-Wing Model\n\n| Wing | b (ft) | c_r (ft) | ε (deg) | x | y | M.A.C. (ft) | Area (sq ft) | Aspect ratio, A |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 1 | 0.175 | 0.499 | 9.93 | 0.18 | 0.08 | 0.333 | 0.0437 | 0.700 |\n| 2 | .323 | .499 | 17.91 | .18 | .08 | .333 | .0805 | 1.292 |\n| 3 | .398 | .493 | 21.96 | .18 | .08 | .329 | .0980 | 1.612 |\n| 4 | .402 | .431 | 25.01 | .18 | .08 | .287 | .0867 | 1.869 |\n| 5 | .409 | .386 | 27.92 | .18 | .08 | .257 | .0790 | 2.114 |\n| 6 | .413 | .360 | 29.84 | .18 | .08 | .240 | .0743 | 2.301 |\n| 7 | .423 | .336 | 32.15 | .18 | .08 | .224 | .0711 | 2.518 |\n| 8 | .433 | .307 | 35.21 | .18 | .08 | .205 | .0665 | 2.812 |\n| 9 | .436 | .279 | 38.01 | .18 | .08 | .186 | .0607 | 3.130 |\n| 10 | .444 | .265 | 39.92 | .18 | .08 | .177 | .0588 | 3.350 |\n| 11 | .463 | .230 | 45.15 | .18 | .08 | .153 | .0532 | 4.023 |\n\n^a Remeasurement shows $y \\approx 0.078$.\n\n(b) Flat-Plate Triangular-Wing Model\n\n| Wing | Sharp leading edge | | | Round leading edge (rad $\\approx$ 0.008 in.) | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | ε (deg) | M.A.C. (ft) | t/c (percent) | ε (deg) | M.A.C. (ft) | t/c (percent) |\n| 1 | 25.13 | 0.289 | 1.3 | 25.00 | 0.283 | 1.3 |\n| 2 | 30.03 | .233 | 1.6 | 30.47 | .226 | 1.7 |\n| 3 | 32.00 | .219 | 1.7 | 31.93 | .206 | 1.8 |\n| 4 | 35.17 | .204 | 1.8 | 35.17 | .200 | 1.9 |\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:37:01.894134+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 43, "total_pages": 62, "image_filename": "19930082918_p43.jpg", "text": "NACA TN 1940\n47\n\n<!-- Image (83, 92, 450, 450) -->\n(a) Unaged.\n\n<!-- Image (512, 92, 900, 451) -->\n(b) Aged 1.0 hour.\n\n<!-- Image (83, 500, 449, 844) -->\n(c) Aged 3.0 hours.\n\n<!-- Image (512, 500, 900, 845) -->\n(d) Aged 10.0 hours.\n\nNACA\nFigure 6.- Effect of aging at $1400^\\circ$ F on microstructure of low-carbon\nN-155 alloy solution-treated 10 hours at $2200^\\circ$ F and water-quenched.\nCross section of bar X1000. Electrolytically etched in 10 percent\nchromic acid.", "timestamp": "2026-07-22T06:37:18.381893+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 22, "total_pages": 33, "image_filename": "19930085900_p22.jpg", "text": "NACA RM L9D20 CONFIDENTIAL 21\n\n[Figure: A photograph of a model in a wind tunnel with water spray. A label on the model reads \"Sta. 34\".]\n\n(a) 8-inch long chines; station 34 to 42; trim, $10^\\circ$\n\n[Figure: A photograph of a model in a wind tunnel with water spray. A label on the model reads \"Sta. 34\". A label in the bottom right corner reads \"NACA L-59856\".]\n\n(b) 32-inch long chines; station 10 to 42; trim, $6.3^\\circ$.\n\nFigure 9.- Effect of length of jet rows at 35 feet per second; $\\frac{1}{4}$-inch-spaced jets.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:37:24.090021+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 19, "total_pages": 29, "image_filename": "19930085899_p19.jpg", "text": "18\nNACA RM No. L9A21\n\n<!-- Image (268, 203, 675, 784) -->\n\nFigure 7.- Concluded.", "timestamp": "2026-07-22T06:37:26.894903+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 2, "total_pages": 92, "image_filename": "19930085951_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:37:27.110018+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 10, "total_pages": 52, "image_filename": "19930085911_p10.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 9\n\nthis altitude. The improvement in combustion in unit A-5 may be attributed to both the change in flame-holder design and the lower values of fuel flow.\n\nIn general, the data indicate that rich fuel-air ratios were detrimental to good combustion. Except for unit A-2, the lowest combustion efficiency was obtained in unit A-4 where the fuel-air ratio (fig. 9(d)) was in excess of 0.084. The combustion efficiency never exceeded 37 percent (fig. 9(e)). Telemeter records showed regions of pulsating, intermittent burning throughout the flight. As a result, some data were not obtainable in unit A-4 between $33\\frac{1}{2}$ and 49 seconds (figs. 9(c) to 9(e)). Unit A-3 also encountered low combustion efficiencies at fuel-air ratios of 0.084 to 0.088 (figs. 8(d) and 8(e)).\n\nThe effect of lean fuel-air ratios on combustion was indicated only by the data for unit A-5. A low combustion efficiency of 40 percent occurred at a fuel-air ratio of 0.043 (figs. 10(d) and 10(e)). As the fuel-air ratio suddenly increased to 0.065, the combustion efficiency increased to 52 percent. In general, for the flight conditions encountered with unit A-5, combustion efficiencies exceeded approximately 40 percent when the ram jet was operating within a fuel-air-ratio range of 0.043 to 0.070.\n\nAn increase in combustion efficiency occurred with an increase in combustion-chamber-inlet static pressure and temperature. For example, the combustion efficiency in unit A-5 (fig. 10(e)) increased from 52 percent at 31 seconds to 91 percent at 38 seconds with increases in combustion-chamber-inlet static pressure from 2800 to 7200 pounds per square foot and static temperature increases from $635^\\circ$ R (fig. 10(d)). This increase in combustion efficiency occurred at approximately constant values of fuel-air ratio (0.062 to 0.065) and combustion-chamber-inlet velocity (150 to 160 ft/sec). During this time interval, however, it was noted that the third set of fuel nozzles became operative and increased the number of nozzles discharging fuel (fig. 10(b)). The resultant change in fuel pattern and degree of fuel atomization may have had some beneficial effect and may partly account for the large increase in combustion efficiency.\n\nThe effect of flame-holder design on the combustion process could be determined only at a fuel-air ratio of 0.067 and low values of combustion-chamber-inlet static pressure of 1000 pounds per square foot and static temperature of $485^\\circ$ R. Under these conditions\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:37:28.647602+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 57, "total_pages": 65, "image_filename": "19930082546_p57.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:37:30.864411+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 37, "total_pages": 149, "image_filename": "19930083192_p37.jpg", "text": "```markdown\nNACA TN 1976\n33\n\nof the same model in two tunnels. Some of the discrepancy may be due to\nthe fact that the gust shape was the same in absolute dimension but\ndiffered slightly on the basis of chords of the two models.\n\n**Effect of power.** - When the slipstream covers the entire span of the\nairplane, the steady-flow slope of the lift curve has been shown to\nincrease about 100 percent for the slipstream corresponding to the climb\ncondition. Since the working velocity has been increased by the intro-\nduction of power, a real increase in the lift-curve slope has probably\nbeen obtained. Under these circumstances, the angle-of-attack change due\nto the gust would be utilized in connection with the increased slope of\nthe lift curve. In actual practice, the suggestion that the power-on\nslope of the lift curve should be used where the slipstream covers most\nof the airplane span is not as serious as it appears at first sight.\nThe highest lift-curve slope is obtained at relatively low forward speed,\na condition at which the load factor due to the gust is small. For\nconventional airplanes, the power-off lift-curve slope appears correct.\n\n**Multiplanes.** - For steady flow conditions the biplane is usually\nrepresented by an equivalent monoplane. This assumption no longer\napplies for the transient lift conditions that exist in a gust. The\nquestion of the proper slope of the lift curve to use for the transient\nconditions is in part answered by the results presented in figure 30.\nOn the basis of the agreement shown in figure 30, the wings should be\nconsidered as acting independently and the slope of the wing lift curve\nas being computed on the basis of the average geometric aspect ratio of\nthe two wings.\n\n**Compressibility.** - The effect of compressibility on the slope of the\nlift curve under unsteady flow conditions is a subject of much interest\nand one for which no experimental evidence is available. It is thought\nthat the lift-curve slope should follow the Glauert factor in the\nsubsonic range, but in the region near the critical Mach number and\nthrough the transonic region the mixed flows combined with transient\nconditions might delay or cancel the effect of compressibility on the\nlift-curve slope. For the time being the usual compressibility factors\ncould be applied up to the critical Mach number and above that, the\nresults of wind-tunnel tests should be utilized.\n\nDownwash\n\nFigure 33 (fig. 9 of reference 31) shows the change in vertical\nvelocity at the tail following a sudden increase in wing circulation for\nseveral values of the distance from the trailing edge of the wing to the\nleading edge of the tail. The change in vertical velocity is plotted as\na function of the distance of the tail surface from the vortex shed by\nthe wing.\n```", "timestamp": "2026-07-22T06:37:31.787425+00:00"}
{"citation_id": "19930086073", "source_url": "https://ntrs.nasa.gov/api/citations/19930086073/downloads/19930086073.pdf", "page_number": 73, "total_pages": 98, "image_filename": "19930086073_p73.jpg", "text": "NACA RM A59E04\n\nLift coefficient, $C_L$\n\nRolling-moment coefficient, $C_l$\n\nYawing-moment coefficient, $C_n$\n\nSide-force coefficient, $C_Y$\n\n$\\delta_{aR}$, deg  \n□ +10.8  \n○ 0  \n◇ -10.8  \n\n(d) $C_L$ vs $C_l$, $C_n$ and $C_Y$.\n\nFigure 15.—Concluded.\n\n71", "timestamp": "2026-07-22T06:37:35.582761+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 2, "total_pages": 22, "image_filename": "19930085928_p2.jpg", "text": "NACA RM No. A9A31 CONFIDENTIAL\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nEXPERIMENTAL INVESTIGATION AT SUPERSONIC SPEEDS OF TWIN-SCOOP\nDUCT INLETS OF EQUAL AREA. IV - SOME EFFECTS OF\nINTERNAL DUCT SHAPE UPON AN INLET ENCLOSING\n37.2 PERCENT OF THE FOREBODY CIRCUMFERENCE\n\nBy Wallace F. Davis, Sherman S. Edwards,\nand George B. Brajnikoff\n\nSUMMARY\n\nTests to determine the recovery of total pressure attainable at\nMach numbers between 1.36 and 2.01 were performed with models having\ntwin-scoop inlets situated on the sides of a long forebody. External\nsupersonic compression occurred through an oblique shock wave created\nby a $12^\\circ$ ramp ahead of an inlet, and boundary-layer removal was\nobtained through slots in the walls of the duct adjacent to the fore-\nbody and extending downstream from the duct entrance. The ducts were\ndesigned to produce supersonic compression in a constricted passage\nbehind the inlet and subsonic diffusion in a channel the shape of\nwhich was calculated to result in local pressure gradients propor-\ntional to the local static pressure. The results of these tests were\ncompared to those of a previous investigation of a model having the\nsame external shape but ducts that expanded from the inlet to a\nconstant diffusion angle at 25 percent of the diffusor length. It\nwas found that the change in internal duct shape caused a large\nincrease in the maximum total-pressure recovery attainable apparently\nbecause the conditions for boundary-layer flow in the diffusor were\nimproved. At Mach numbers of 1.7 and less, the pressure recovery\nwas within two percent of that associated with nose inlets.\n\nINTRODUCTION\n\nThe results of the investigation described in reference 1 show\nthat the recovery of total pressure attained with a twin-scoop inlet\nin the presence of a boundary layer at Mach numbers between 1.36\nand 2.01 was very nearly equal to that of a normal shock wave\noccurring at the free-stream Mach number. In order to attain this\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:37:37.285495+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 71, "total_pages": 114, "image_filename": "19930086061_p71.jpg", "text": "NACA RM L9J07\n67\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\nP\n-3\n-2\n-1\n0\n1\n-3\n-2\n-1\n0\n1\n20°\n(c) $\\psi = 20^\\circ$.\n\nLeft semispan\nRight semispan\nUpper\nLower\nP\nP\n-3\n-2\n-1\n0\n1\n-3\n-2\n-1\n0\n1\n35°\n(d) $\\psi = 35^\\circ$.\nNACA\nFigure 22.- Concluded.", "timestamp": "2026-07-22T06:37:40.038111+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 21, "total_pages": 26, "image_filename": "19930085890_p21.jpg", "text": "```markdown\n20\nNACA RM No. E9C11\n\nTheoretical; equilibrium\nexpansion (reference 1)\nTheoretical; equilibrium\nexpansion; 95-percent\ndiborane; total nozzle\ndivergent angle, 30°\nExperimental, corrected\nExperimental\n\nVolume specific impulse, $I_d$, $\\frac{\\text{lb-sec}}{\\text{cu ft}} \\times \\frac{1}{62.4}$\n\nRatio of fuel weight to total propellant weight\n\n[Figure: Graph showing volume specific impulse vs. ratio of fuel weight to total propellant weight. The graph contains four curves: two dashed lines representing theoretical equilibrium expansion (one for reference 1, one for 95-percent diborane with a 30° nozzle angle), one dash-dot line for corrected experimental data, and one solid line with circular data points for uncorrected experimental data. The y-axis ranges from 100 to 240, and the x-axis ranges from .20 to .60. A NACA logo is present in the bottom right corner of the plot area.]\n\nFigure 6. - Theoretical and experimental volume specific impulse\nin 100-pound-thrust rocket engine using liquid diborane and\nliquid oxygen.\n```", "timestamp": "2026-07-22T06:37:42.179709+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 44, "total_pages": 62, "image_filename": "19930082918_p44.jpg", "text": "NACA TN 1940\n49\n\n[Figure: Micrograph showing grain structure with some precipitates and grain boundaries]\n(e) Aged 30 hours.\n\n[Figure: Micrograph showing grain structure with more numerous and finer precipitates]\n(f) Aged 100 hours.\n\n[Figure: Micrograph showing grain structure with a very high density of fine precipitates]\n(g) Aged 1000 hours.\n\nFigure 6.- Concluded.\nNACA", "timestamp": "2026-07-22T06:37:49.033325+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 23, "total_pages": 33, "image_filename": "19930085900_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:37:49.494931+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 6, "total_pages": 47, "image_filename": "19930085919_p6.jpg", "text": "NACA RM No. A9C21 CONFIDENTIAL 5\n\nMODEL AND TESTS\n\nThe semispan wing used for this investigation had its leading edge swept back 63°, an aspect ratio of 3.5 based on the geometry of the complete wing, a taper ratio (ratio of tip chord to root chord) of 0.25, no twist, no dihedral, and the NACA 64A006 profile parallel to the plane of symmetry. The model wing is shown in figure 1 mounted from the floor of the Ames 7- by 10-foot wind tunnel No. 2. Model dimensions are presented in figure 2.\n\nA gap of one-eighth inch existed between the turntable and the extension of the wing spar which passed through the turntable to support the model. The clearance between the tunnel floor and the model was about one-quarter inch except near the nose of the long fuselage where the gap was about three-quarters inch.\n\nThe fuselage used in part of this investigation was semi-circular in cross section and had a fineness ratio of 12.5. This fuselage is hereafter referred to as the long fuselage.¹ Due to possible effects of the wind-tunnel walls on the experimental results, the maximum angle of attack employed with this fuselage was 26°. To allow for a greater angle-of-attack range for the major portion of the investigation, this fuselage was shortened to a fineness ratio of 10.5. This fuselage is referred to as the short fuselage.² The wing is shown in combination with the long and short fuselages in figure 3.\n\nThe model was tested with a 0.25-chord split flap in four chordwise positions on the wing with hinges lines along lines of constant-percent wing chord (40, 60, 75, and 100 percent of the\n\n¹Equation for contour of long fuselage (see fig. 2)\n\n$$\nr = 0.680 \\left[ 1.000 - \\left( 1.000 - \\frac{x}{8.500} \\right)^2 \\right]^{3/4}\n$$\n\n²Equations for contour of short fuselage (see fig. 2)\n\nNose: $ r = \\sqrt{77.371 - (x - 3.116)^2} - 8.226 $\n\nTail: $ r = \\sqrt{0.918 - (x - 0.718)^2} - 0.635 $\n\nThe ordinates of the center portion of the short fuselage were identical to the ordinates of the center portion of the long fuselage from 51.00 inches to 183.60 inches from the nose of the long fuselage.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:37:53.547543+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 58, "total_pages": 65, "image_filename": "19930082546_p58.jpg", "text": "NACA TN No. 1870\n57\n\n<!-- Image (169, 156, 831, 937) -->\n\nFigure 19.- Charts for estimating the maximum free-space pressures near the plane of rotation of a rotating propeller.", "timestamp": "2026-07-22T06:38:00.377382+00:00"}
{"citation_id": "19930085899", "source_url": "https://ntrs.nasa.gov/api/citations/19930085899/downloads/19930085899.pdf", "page_number": 20, "total_pages": 29, "image_filename": "19930085899_p20.jpg", "text": "```markdown\nNACA RM No. L9J21\n\nAngle of attack, $\\alpha$, deg\nLift coefficient, $C_L$\nM\n1.15 $\\triangle$ 0\n1.10 $\\nabla$ 0\n1.08 $\\triangle$ 0\n1.05 $\\nabla$ 0\n1.03 $\\triangle$ 0\n1.00 $\\square$ 0\n.98 $\\triangle$ 0\n.95 $\\nabla$ 0\n.93 $\\triangle$ 0\n.90 $\\nabla$ 0\n.88 $\\nabla$ 0\n.85 $\\nabla$ 0\n.80 $\\square$ 0\n.70 $\\square$ 0\n.60 $\\circ$ 0\n\nPitching-moment coefficient, $C_m$\nLift coefficient, $C_L$\nM\n1.15 $\\triangle$ 0\n1.10 $\\nabla$ 0\n1.08 $\\triangle$ 0\n1.05 $\\nabla$ 0\n1.03 $\\triangle$ 0\n1.00 $\\square$ 0\n.98 $\\triangle$ 0\n.95 $\\nabla$ 0\n.93 $\\triangle$ 0\n.90 $\\nabla$ 0\n.88 $\\nabla$ 0\n.85 $\\nabla$ 0\n.80 $\\square$ 0\n.70 $\\square$ 0\n.60 $\\circ$ 0\n\nDrag coefficient, $C_D$\nLift coefficient, $C_L$\n[Figure: NACA logo]\n\nFigure 8.— Wing-fuselage aerodynamic characteristics for model with 45° sweptback wing, aspect ratio 4, taper ratio 0.6, and NACA 65A006 airfoil.\n\n19\n```", "timestamp": "2026-07-22T06:38:01.347886+00:00"}
{"citation_id": "19930083192", "source_url": "https://ntrs.nasa.gov/api/citations/19930083192/downloads/19930083192.pdf", "page_number": 38, "total_pages": 149, "image_filename": "19930083192_p38.jpg", "text": "34\nNACA TN 1976\n\nFigure 33 indicates that the tail surface first encounters a gradual increase in upwash until the leading edge crosses the center of the shed vortex, when, following a violent change in direction, the downwash approaches its steady-state value. The curves for different tail lengths indicate quite clearly that the effect of increasing the tail length is to increase the time delay of the velocity change in proportion to the tail length. The violence of the direction change as the tail penetrates the vortex depends on the vertical location of the tail relative to the vortex, and, until more accurate predictions can be made of the location of the vortex, such changes might be disregarded. Figure 33 indicates that the assumption proposed in references 31 and 32 of a space lag between the change in lift on the wing and the change in angle of attack at the tail is reasonable and should be taken into account in any detailed calculations.\n\nThe indications to date are that the results obtained by utilizing a space lag of downwash have been in fair agreement with experiment, but the corresponding calculations neglecting the space lag of downwash have not been made to determine the seriousness of the error of neglecting this quantity.\n\n### Maximum Lift Coefficient\n\nReference 19 shows that for an airplane model traversing a sharp-edge gust at its steady-flow maximum lift coefficient, the maximum lift coefficient during the action of the gust was not limited to the steady-flow value. Farren in reference 33 made tests of two-dimensional airfoils at constant angular velocities through maximum lift and return. Angle-of-attack variations range as high as $12\\frac{1}{2}^\\circ$ per chord of travel and the results indicated that the maximum lift coefficient could range from 30 to 50 percent above the steady-flow value.\n\nNo concise estimates of maximum lift coefficient during sudden changes in angle of attack can be made, but the sketchy information that is available indicates that a conservative estimate for incompressible-flow conditions is an increase of about 25 percent over the steady-flow value. In the transonic region, the effects of compressibility may limit the maximum lift coefficient, and in this range the investigations made in the various wind tunnels would be pertinent.\n\n### RIGID-BODY REACTIONS\n\nThe available data on gust structure and aerodynamics have been analyzed and the next problem is the determination of the behavior of", "timestamp": "2026-07-22T06:38:03.039635+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 11, "total_pages": 52, "image_filename": "19930085911_p11.jpg", "text": "10 CONFIDENTIAL NACA RM E9F22\n\nwith the ducted-type flame holder, a combustion efficiency of 20 percent was obtained at 22 seconds (fig. 8(e)) as compared with a combustion efficiency of 46 percent obtained at 16.5 seconds with the rake-type flame holder (fig. 10(e)).\n\nThe performance of unit A-5 (fig. 10) indicated that a high combustion efficiency of 91 percent could be obtained at a free-stream Mach number of 1.70, which sustained a diffuser total-pressure recovery of approximately 0.90. A maximum net acceleration of 2.0 g's and a thrust coefficient of 0.56 were produced. A gas total-temperature ratio of 5.1 existed, which was equivalent to an exhaust-gas total temperature of 4050° R.\n\nDiffuser Total-Pressure Recovery\n\nThe total-pressure recovery across the diffuser is shown in figure 12 as a function of free-stream Mach number for all four ram-jet units. Lines of constant gas total-temperature ratio were faired according to the collective data points.\n\nAt a constant value of free-stream Mach number, a decrease in gas total-temperature ratio was accompanied by a decrease in total-pressure recovery largely due to increasing shock losses within the diffuser. For example, at a free-stream Mach number of 1.20, the diffuser total-pressure recovery decreased from 0.94 to 0.52 with a decrease in gas total-temperature ratio from 5.0 to 1.0. The decrease in gas total-temperature ratio resulted in diffuser-outlet conditions of higher velocity, lower static pressure, and a reduction in total pressure necessary to maintain mass continuity. When supersonic velocity existed within the diffuser, this reduction in total pressure was made possible only by the presence of a normal shock with its accompanying total-pressure loss. The minimum value of free-stream Mach number at which shock occurred within the diffuser was approximately 0.60 at a gas total-temperature ratio of 1.2 (fig. 7) at 16 seconds after release with the transition from subsonic to supersonic flow at station 2 ($4\\frac{5}{8}$ in. downstream of the diffuser inlet).\n\nThrust Coefficient\n\nThe effects of free-stream Mach number and gas total-temperature ratio on the net-thrust coefficients are shown in figure 13. As defined in the appendix, the net thrust is the total change in momentum of the air and the fuel flowing through the ram jet. At constant gas total-temperature ratios, the thrust coefficient increased with increasing free-stream Mach numbers until the total-pressure recovery declined rapidly because of shock in the diffuser.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:03.504706+00:00"}
{"citation_id": "19930085951", "source_url": "https://ntrs.nasa.gov/api/citations/19930085951/downloads/19930085951.pdf", "page_number": 3, "total_pages": 92, "image_filename": "19930085951_p3.jpg", "text": "NACA RM L9D29\n\n[annotation: CONFIDENTIAL]\n\nClassification Changed to\nUNCLASSIFIED\n\nNACA Research Abstract #21,\ndated April 24, 1952 - 3/28/52\n\nDate\n6/23/52\n\nBy\nD. E. Rowland\nNACA\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nTHE EFFECT OF BLADE-SECTION THICKNESS RATIOS\nON THE AERODYNAMIC CHARACTERISTICS\nOF RELATED FULL-SCALE PROPELLERS\nAT MACH NUMBERS UP TO 0.65\n\nBy Julian D. Maynard and Seymour Steinberg\n\nSUMMARY\n\nThe results of an investigation of two full-scale NACA propellers are presented for a range of blade angles from 20° to 55° at airspeeds up to 500 miles per hour. These results are compared with the results from previous investigations of five related NACA propellers to evaluate the effects of blade-section thickness ratios on propeller aerodynamic characteristics.\n\nThe envelope efficiencies of all the NACA propellers are high at the lower rotational speeds at which the adverse effects of compressibility are small. The highest efficiencies, about 93 percent at a helical-tip Mach number of 0.9 and 84 percent at a helical-tip Mach number of 1.1, reflect the importance of using thin, efficient airfoil sections throughout the blade. For propeller operation at constant rotational speed and power at helical-tip Mach numbers below 0.8 a reduction in blade-section thickness from 12 to 8 percent at the 0.7 radius, or approximately one-third all along the radius, results in gains in propeller efficiency up to 10 percent.\n\nThe maximum efficiency of a propeller operating at a helical-tip Mach number of 1.1 and air-stream Mach number of 0.625 may be increased approximately 20 percent by reducing the blade-section thickness from 12 to 5 percent at the 0.7 radius. At this same condition of operation for propellers having blade-section thicknesses between 12 and 8 percent at the 0.7 radius the maximum efficiency increases approximately 3 percent for each decrease in thickness of 1 percent at the 0.7 radius. For blade-section thicknesses between 8 and 5 percent at the 0.7 radius the rate of increase in propeller efficiency with reductions in blade-section thickness is smaller, but further reductions in thickness may still improve the maximum efficiency of propellers operating at high forward speeds with helical-tip Mach numbers as high as 1.1.\n\n[annotation: UNCLASSIFIED]\n[annotation: CONFIDENTIAL]", "timestamp": "2026-07-22T06:38:05.543618+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 22, "total_pages": 26, "image_filename": "19930085890_p22.jpg", "text": "NACA RM NO. E9C11\n\n(a) Engine chamber, downstream.\nNACA\nC-21505\n5-20-48\n\n(b) Exhaust nozzle, upstream.\nNACA\nC-21676\n6-10-48\n\n(c) Injection head and nozzle.\nNACA\nC-21679\n6-10-48\n\n(d) Spacer.\nNACA\nC-21678\n6-10-48\n\nFigure 7. - Deposits on inner walls of rocket engine.\n\n21", "timestamp": "2026-07-22T06:38:06.512891+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 72, "total_pages": 114, "image_filename": "19930086061_p72.jpg", "text": "68\nNACA RM L9J07\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\n-1 P\n0\n1\nRight semispan\n\nUpper\nLower\n\n(a) $\\psi = 0^\\circ$\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n\n-3\n-2\n-1 P\n0\n1\nRight semispan\n\nUpper\nLower\n\n10°\n\n(b) $\\psi = 10^\\circ$\n\nNACA\n\nFigure 23.- Pressure distribution about wing 3 at various angles of yaw; $\\alpha = 14.1^\\circ$.", "timestamp": "2026-07-22T06:38:06.718691+00:00"}
{"citation_id": "19930085928", "source_url": "https://ntrs.nasa.gov/api/citations/19930085928/downloads/19930085928.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930085928_p3.jpg", "text": "2\nCONFIDENTIAL\nNACA RM No. A9A31\n\nrecovery, three design features were found to be necessary: (1) The scoops had to enclose a relatively small portion of the forebody circumference so that the proportion of boundary layer to unimpeded air flowing into the diffusor was small; (2) the intake Mach number had to be reduced by external compression through an oblique shock wave; and (3) some of the boundary layer that flowed into the scoops had to be forced out of the diffusor through slots in the duct walls immediately behind the inlet. The tests showed that if the intake Mach number were reduced by deflecting the stream with a ramp ahead of the inlet to create an oblique shock wave, ramp angles greater than about $12^\\circ$ caused no additional compression. This limit existed because the boundary layer thickened ahead of the break in the surface when greater ramp angles were used; the boundary layer filled the break and thereby maintained an effective deflection angle of $12^\\circ$. The slots in the duct walls apparently improved the flow in the subsonic diffusor by reducing the amount of retarded air and delaying separation until the flow was more fully diffused.\n\nSince the ramp and the slots produced a large, though limited, improvement in the pressure recovery attainable with this inlet, it was reasoned that additional methods for creating supersonic compression and improvements in the boundary-layer flow might further increase the recovery. In an attempt to produce supersonic compression besides that through the oblique shock wave from the ramp, a convergent passage was added immediately downstream of the duct entrances of the configuration described in reference 1. This passage was intended to produce nearly isentropic compression of the flow from the intake Mach number to a lower supersonic Mach number at the throat of the duct. The effect of this additional supersonic compression should be a reduction in the pressure losses due to the shock waves through which the flow is decelerated to subsonic speed. In order to improve the flow in the divergent subsonic diffusor beyond the improvement caused by the slots, the shape of the duct downstream of the throat was changed to decrease the adverse pressure gradient in the high-velocity section and so to delay separation of the boundary layer. The present report describes the results of tests of models having these additional considerations in the design of the internal ducts.\n\nSYMBOLS\n\nA area\n\nH total pressure\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:07.707291+00:00"}
{"citation_id": "19930085862", "source_url": "https://ntrs.nasa.gov/api/citations/19930085862/downloads/19930085862.pdf", "page_number": 35, "total_pages": 60, "image_filename": "19930085862_p35.jpg", "text": "NACA RM No. L9A07\n33\n\n$C_{Na}$\n$\\delta_a$ (deg)\n-25\n-20\n-15\n-10\n-5\n0\n5\n10\n15\n20\n25\n\n$C_n$\n\n$C_l$\n$\\alpha$, deg\n\n(a) $C_l$, $C_n$, and $C_{Na}$ against $\\alpha$.\n\nFigure 9.— Aileron characteristics of wing with extensible leading-edge and split flaps and fences.", "timestamp": "2026-07-22T06:38:08.332453+00:00"}
{"citation_id": "19930083221", "source_url": "https://ntrs.nasa.gov/api/citations/19930083221/downloads/19930083221.pdf", "page_number": 41, "total_pages": 47, "image_filename": "19930083221_p41.jpg", "text": "NACA TN No. 1824\n39\n\n$$\n\\frac{C_L}{\\alpha \\tan \\theta} = \\frac{\\pi}{2} \\frac{A}{\\tan \\theta} \\left( 1 - \\frac{t_o^2}{s_o^2} \\right)\n$$\n\n[Figure: Two diagrams showing pressure distributions ($\\Delta p / q \\alpha$) on wing surfaces with axes $x$ and $y$.]\n\nFigure 18. Pressure distributions for triangular and swept-back wings at $M_o = 1$.\n\nand\n\n$$\n\\frac{C_{D_i}}{\\alpha^2 \\tan \\theta} = \\frac{A}{\\tan \\theta} \\left[ \\frac{k_3^2 \\pi}{4} - \\frac{E_3' - k_3^2 K_3'}{(s_o/c_o \\tan \\theta)} \\right] \\quad (73)\n$$\n\nwhere\n\n$$\nk_3' = \\frac{t_o}{s_o} = \\sqrt{1 - k_3^2}\n$$\n\nThese coefficients are plotted as a function of $A/\\tan \\theta$ in figure 19. It is shown that the values of $C_{D_i}/\\alpha^2 \\tan \\theta$ and $C_{L_i}/\\alpha \\tan \\theta$ for finite aspect ratio swept-back wings are always less than the corresponding values for the triangular wing ($A/\\tan \\theta = 4$).", "timestamp": "2026-07-22T06:38:09.353301+00:00"}
{"citation_id": "19930082483", "source_url": "https://ntrs.nasa.gov/api/citations/19930082483/downloads/19930082483.pdf", "page_number": 60, "total_pages": 78, "image_filename": "19930082483_p60.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:38:10.170781+00:00"}
{"citation_id": "19930085913", "source_url": "https://ntrs.nasa.gov/api/citations/19930085913/downloads/19930085913.pdf", "page_number": 9, "total_pages": 34, "image_filename": "19930085913_p9.jpg", "text": "8\nNACA RM L9F24\n\nTEST PROCEDURE\n\nSince flutter is usually destructive, recognition of flutter,\nrecording of the necessary data, and reduction of the airspeed to save\nthe model must be accomplished in a very short time. The airspeed was\nincreased slowly, and at the flutter point, oscillograph records were\ntaken and the tunnel conditions were recorded. The first three natural\nfrequencies of each model at zero airspeed for the various weight\npositions were recorded before the model was flutter-tested. After\neach model had been made to flutter with various weighted conditions,\nit was retested with no weight to establish whether or not it had been\ndamaged by flutter. In addition, the nodal-line patterns associated\nwith the second and third natural frequencies of the models at zero\nairspeed (fig. 2) were obtained.\n\nPRESENTATION OF RESULTS\n\nExperimental results, obtained from the flutter tests of wings\nwith sweepback angles of $0^\\circ$, $45^\\circ$, and $60^\\circ$ and carrying a single concen-\ntrated weight at a series of spanwise and two chordwise positions, are\npresented in table I and in figures 3 and 4. In table I the quantities\nlisted are dynamic pressure, flutter velocity, Mach number, natural and\nflutter frequencies, and phase-angle relationships of the bending and\ntorsional stresses for the corresponding second and third natural and\nflutter frequencies. A sketch of each model configuration is included\nin the table with its corresponding data.\n\nThe oscillograph records taken at flutter for the 95 flutter tests\nare shown in figure 1. The gage traces in the records are numbered\nfrom left to right and are: (1) root torsion, (2) root bending,\n(3) tip torsion, and (4) tip bending. The gage traces are marked at\nthe top of each record with their appropriate attenuations. The run\nnumbers are given in the lower left-hand corner of each record.\n\nThe second and third natural-frequency nodal lines of each model\nconfiguration weighted at the leading edge are shown in figure 2. The\nprogressive change in these nodal lines with spanwise weight position\nis illustrated.\n\nIn figure 3 the first three natural frequencies and the flutter\nfrequency are plotted against spanwise weight position for each sweep\nangle and chordwise weight position. These plots show the relation\nbetween the flutter frequency and the first three natural frequencies\nof the wing for a given weight position.", "timestamp": "2026-07-22T06:38:13.768882+00:00"}
{"citation_id": "19930085922", "source_url": "https://ntrs.nasa.gov/api/citations/19930085922/downloads/19930085922.pdf", "page_number": 7, "total_pages": 22, "image_filename": "19930085922_p7.jpg", "text": "6\nNACA RM No. L9C23\n\n3. Vertical fins, mounted at approximately the midspan section of each wing panel, had little effect on the damping-in-roll characteristics of the wing but increased the effectiveness of the ailerons in producing rolling moment.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va.\n\nREFERENCES\n\n1. Donlan, Charles J., and Kuhn, Richard E.: Estimated Transonic Flying Qualities of a Tailless Airplane Based on a Model Investigation. NACA RM No. L9D08, 1949.\n\n2. Herriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnels, with Consideration of the Effect of Compressibility. NACA RM No. ATB28, 1947.\n\n3. Bird, John D.: Some Theoretical Low-Speed Span Loading Characteristics of Swept Wings in Roll and Sideslip. NACA TN No. 1839, 1949.\n\n4. Queijo, M. J., and Lichtenstein, Jacob H.: The Effects of High-Lift Devices on the Low-Speed Stability Characteristics of a Tapered $37.5^\\circ$ Sweptback Wing of Aspect Ratio 3 in Straight and Rolling Flow. NACA RM No. L8I03, 1948.\n\n5. MacLachlan, Robert, and Letko, William: Correlation by Two Experimental Methods of Determining the Rolling Characteristics of Unswept Wings. NACA TN No. 1309, 1947.", "timestamp": "2026-07-22T06:38:14.000594+00:00"}
{"citation_id": "19930082918", "source_url": "https://ntrs.nasa.gov/api/citations/19930082918/downloads/19930082918.pdf", "page_number": 45, "total_pages": 62, "image_filename": "19930082918_p45.jpg", "text": "NACA TN 1940\n51\n\n[Figure: (a) Unaged.]\n\n[Figure: (b) Aged 1.0 hour.]\n\n[Figure: (c) Aged 10 hours.]\n\nFigure 7.- Electron micrographs of replicas (X8500) prepared from low-carbon N-155 alloy solution-treated 10 hours at 2200° F, water-quenched, and aged at 1400° F.\n\nNACA", "timestamp": "2026-07-22T06:38:16.065060+00:00"}
{"citation_id": "19930082546", "source_url": "https://ntrs.nasa.gov/api/citations/19930082546/downloads/19930082546.pdf", "page_number": 59, "total_pages": 65, "image_filename": "19930082546_p59.jpg", "text": "58\nNACA TN No. 1870\n\n<!-- Image (159, 131, 826, 903) -->\n\nFigure 19.— Continued.\n(c) mB = 4.\n(a) mB = 5.", "timestamp": "2026-07-22T06:38:16.834312+00:00"}
{"citation_id": "19930085900", "source_url": "https://ntrs.nasa.gov/api/citations/19930085900/downloads/19930085900.pdf", "page_number": 24, "total_pages": 33, "image_filename": "19930085900_p24.jpg", "text": "```markdown\nNACA RM L9D20\n23\n\nCONFIDENTIAL\n\n<!-- Image (169, 125, 858, 888) -->\n\nFigure 10.- Comparison of slanted and straight jets; $\\frac{1}{4}$-inch-spaced jets;\nstation 10 to 42.\nCONFIDENTIAL\n```", "timestamp": "2026-07-22T06:38:17.262878+00:00"}
{"citation_id": "19930085491", "source_url": "https://ntrs.nasa.gov/api/citations/19930085491/downloads/19930085491.pdf", "page_number": 42, "total_pages": 72, "image_filename": "19930085491_p42.jpg", "text": "NACA RM No. A8J04 CONFIDENTIAL 41\n\n[Figure: Exploded view of model components including two wings, a central fuselage section, and a nose cone assembly. A scale bar labeled \"INCHES\" with markings for 1 and 2 inches is shown. Annotation: NACA A-11962-A]\n\nFigure 2.— Exploded view of model.\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:18.093610+00:00"}
{"citation_id": "19930085890", "source_url": "https://ntrs.nasa.gov/api/citations/19930085890/downloads/19930085890.pdf", "page_number": 23, "total_pages": 26, "image_filename": "19930085890_p23.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T06:38:23.357931+00:00"}
{"citation_id": "19930086061", "source_url": "https://ntrs.nasa.gov/api/citations/19930086061/downloads/19930086061.pdf", "page_number": 73, "total_pages": 114, "image_filename": "19930086061_p73.jpg", "text": "NACA RM L9J07\n69\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n-3\n-2\nP -1\n0\n1\nRight semispan\nUpper\nLower\n\n20°\n\n(c) $\\Psi = 20^\\circ$\n\n-3\n-2\n-1 P\n0\n1\nLeft semispan\n-3\n-2\nP -1\n0\n1\nRight semispan\nUpper\nLower\n\n35°\n\n(d) $\\Psi = 35^\\circ$\n\nFigure 23.- Concluded.\nNACA", "timestamp": "2026-07-22T06:38:26.466543+00:00"}
{"citation_id": "19930085911", "source_url": "https://ntrs.nasa.gov/api/citations/19930085911/downloads/19930085911.pdf", "page_number": 12, "total_pages": 52, "image_filename": "19930085911_p12.jpg", "text": "NACA RM E9F22 CONFIDENTIAL 11\n\nAt sonic free-stream velocity, a gas total-temperature ratio of at least 2.0 was necessary for a positive thrust coefficient. At this velocity, the thrust coefficient varied from -0.20 to 0.48 with variations in the gas total-temperature ratio from 1.0 to 6.0.\n\nExternal Drag Coefficient\n\nThe effect of free-stream Mach number and gas total-temperature ratio on the external drag coefficient is shown in figure 14. The external drag equals the total change of momentum of the air flowing outside the ram jet and therefore includes the additive drag at the diffuser inlet as well as the total external drag on the shell and the fins. The term additive drag is more fully discussed in reference 3. The dashed curve represents the minimum external drag coefficients encountered at various free-stream Mach numbers. These minimum values occurred at conditions of minimum additive drag and maximum engine air flow when the external-flow conditions ahead of the diffuser inlet were unaffected by variations in heat addition. Increasing the heat addition above a certain value will reduce the air flow; consequently, the divergence of the streamlines ahead of the inlet becomes greater. Increased pressures acting on these diverging streamlines give rise to increased additive drag. The increased pressures also increase the wave drag on the shell exterior. The effect of an increase in heat addition on the external drag coefficient was most noticeable at supersonic Mach numbers. For example, at a free-stream Mach number of 1.70, increasing the gas total-temperature ratio from 4.0 to 5.0 increased the external drag coefficient from 0.17 to 0.32.\n\nThe lowest value of external drag coefficient (approximately 0.115) occurred at free-stream Mach numbers of 0.90 to 1.00. The maximum external drag coefficients for any given gas total-temperature ratio occurred at free-stream Mach numbers between 1.10 and 1.20. When the ram jet was operating at the design condition (gas total-temperature ratio of 4.0 and free-stream Mach number of 1.60), the external drag coefficient was 0.169, which was the minimum value for the free-stream Mach number of 1.60.\n\nSUMMARY OF RESULTS\n\nFrom the data obtained from free-flight investigations of four 16-inch-diameter supersonic ram-jet units with a range of free-stream Mach numbers of 0.38 to 1.73 and gas total-temperature ratios between 1.0 and 6.6, the following results were observed:\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:27.315865+00:00"}
{"citation_id": "19930085919", "source_url": "https://ntrs.nasa.gov/api/citations/19930085919/downloads/19930085919.pdf", "page_number": 7, "total_pages": 47, "image_filename": "19930085919_p7.jpg", "text": "6 CONFIDENTIAL NACA RM No. A9C21\n\nwing chord), and in two chordwise positions with hinge lines normal to the air stream. The model was also investigated with a split flap of triangular plan form with its hinge line coincident with the trailing edge of the wing. All of the split flaps had the same area and extended from the fuselage to the midsemispan of the wing. The dimensions and positions of the split flaps on the wing are shown in figure 4.\n\nThe model was investigated with an elevon having chords equal to 25 percent of the local wing chord, and with an elevon of constant chord. Both elevons extended from the midsemispan to the wing tip and had unsealed radius noses. The dimensions of the elevons are given in figure 2.\n\nSectional views of the leading-edge flaps and the sharp leading edge are shown in figure 5. The model was tested with these devices having span equal to 50 and 100 percent of the wing span. The 50-percent-span leading-edge flaps extended from the midsemispan to the wing tip; whereas the 50-percent-span sharp leading edge extended from the midsemispan to the wing-fuselage juncture. Photographs of the model with the full-span drooped-nose and extended-nose flaps are shown in figure 6.\n\nMost of the tests were conducted at a dynamic pressure of 50 pounds per square foot which corresponded to a Reynolds number of 4.2 million. However, to investigate possible dynamic scale effects, some tests were performed throughout a Reynolds number range of 2.5 to 7.2 million.\n\nRESULTS AND DISCUSSION\n\nPlain Wing and Wing-Fuselage Combinations\n\nThe results of tests of the plain wing and wing-fuselage combinations are presented in figure 7. The following characteristics of the plain wing are noted just above a lift coefficient of 0.2: (1) The rate of change of lift with angle of attack increased, and (2) the aerodynamic center shifted rearward. Observations of the flow in the boundary layer, by means of short tufts of thread attached to the wing surface, indicated that a local region of flow separation occurred near the wing leading edge in the vicinity of the wing tip at a lift coefficient of approximately 0.2. The following characteristics of the plain wing are noted in figure 7 just above a lift coefficient of 0.4: (1) The rate of change of lift with angle\n\nCONFIDENTIAL", "timestamp": "2026-07-22T06:38:28.978775+00:00"}