PDF/ocr_dataset_1024.jsonl

{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 18, "total_pages": 53, "image_filename": "19930082542_p18.jpg", "text": "NACA TN No. 1867\n\nhot-cold-work probably results from precipitation during testing which strengthens and maintains the high level of strength developed by the strain hardening. (4) Hot-worked materials have varied properties because all the trends discussed under the first three items could be influencing the properties.\n\nCONCLUSIONS\n\nThe properties of bar stock from one heat of low-carbon N-155 alloy, as measured by tensile tests at room temperature and rupture tests at $1200^\\circ$ F, have been systematically measured over wide limits after solution treatments, aging treatments, and hot-cold-work.\n\nYield strengths at room temperature ranging from 30,000 to 134,000 psi at 0.02-percent offset can be produced in the alloy. The same treatments will result in rupture strengths for fracture in 100 hours at $1200^\\circ$ F from 40,000 to 66,000 psi and for fracture in 1000 hours from 35,000 to 56,000 psi. These rupture strength ranges are equivalent to such extreme differences in actual time for fracture as 100 to 600,000 hours under a stress of 40,000 psi for the same alloy with different treatments.\n\nOn the basis of the yield strength for 0.02-percent offset at room temperature and rupture properties at $1200^\\circ$ F the following standard-type treatments are the best for the alloy:\n\n1. For solution-treating the alloy only, the optimum temperature is about $2100^\\circ$ F. The yield strength will be below 40,000 psi and the rupture strengths for 100 and 1000 hours will be about 51,000 and 40,000 psi. Treatments at lower and higher temperatures will result in lower rupture strengths. Higher temperatures also will cause excessively low ductility and brittleness at stress concentrations during rupture testing.\n\n2. The best aging treatment, $1350^\\circ$ to $1400^\\circ$ F for either 2 or 24 hours, will result in yield strengths between 40,000 and 50,000 psi and rupture strengths of about 50,000 psi for 100 hours and 35,000 to 43,000 psi for 1000 hours. Major effects of aging treatments will be increased ductility in the rupture test and low rupture strength for time periods longer than 1000 hours.\n\n3. High strength is dependent on hot-cold-work. Maximum properties will be developed by 10- to 15-percent reduction at temperatures below $1400^\\circ$ F. The solution temperature prior to hot-cold-working should be between approximately $1950^\\circ$ and $2100^\\circ$ F. Yield strengths between 90,000 and 110,000 psi and rupture strengths for fracture in 100 and 1000 hours of better than 60,000 and 52,000 psi will be obtained. These treatments also produce the highest rupture strength at prolonged time periods.", "timestamp": "2026-07-22T04:57:19.376806+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 16, "total_pages": 78, "image_filename": "19930082618_p16.jpg", "text": "14 NACA TN 1945\n\nStalling characteristics.- In airplane design problems the manner in which the airfoil stalls is frequently of great importance. The lift data in parts (a) and (b) of figures 1 to 15 show that the type of stall depends upon the Reynolds number, airfoil design, surface condition, and whether a split flap is employed. In general, the data for the plain, smooth NACA 6-series sections show that the stall becomes less abrupt as the airfoil thickness and camber are increased and as the Reynolds number is reduced. The favorable effect of a decreasing Reynolds number on the character of the stall is not evident in the data for the NACA 6-series airfoils of 15- and 18-percent thickness and for the airfoil of 0.6 design lift coefficient. These sections, however, show favorable stalling characteristics at all Reynolds numbers as compared with the rather abrupt stalls shown by the thinner sections and sections of smaller camber at the higher Reynolds numbers. Variations in thickness form corresponding to positions of minimum pressure on the basic thickness form at zero lift from 30 to 60 percent chord do not appear to have any effect upon the character of the stall of the 15-percent-thick, smooth airfoil sections. Rearward movement of the minimum-pressure point may, however, have some small adverse effect upon the stalling characteristics of airfoils thinner than 15 percent of the chord as indicated by data corresponding to Reynolds numbers from $3.0 \\times 10^6$ to $9.0 \\times 10^6$ for 12-percent-thick airfoils having different positions of minimum pressure (reference 1).\n\nIn the smooth surface condition the two NACA 230-series sections are seen to possess extremely undesirable stalling characteristics at nearly all Reynolds numbers, whereas both of the NACA 44-series sections have very good stall characteristics throughout the Reynolds number range investigated. The stall of the NACA 0012 section is very acute at the higher Reynolds numbers but, like the NACA $64_1$-012 airfoil, reductions in the Reynolds number have a somewhat favorable effect.\n\nIn the rough surface condition, nearly all of the plain airfoils have good stalling characteristics at most Reynolds numbers. The NACA 230-series sections, and at the higher Reynolds numbers the NACA 0012 section, are notable exceptions, for even in the rough condition the stalling characteristics of these airfoils are rather undesirable at most Reynolds numbers.\n\nWith 0.20c split flaps deflected $60^\\circ$, the stalling characteristics of the smooth NACA 6-series sections do not vary in an entirely consistent manner with either Reynolds number or airfoil design. In some cases, decreasing the Reynolds number improves the stalling characteristics (examples, NACA $64$-409 and NACA $64_1$-412 sections); in other cases, decreasing the Reynolds number does not seem to affect the stall (NACA $64_3$-418); whereas in still other cases, the stall seems to be affected slightly adversely by reducing the Reynolds number (NACA $63_2$-415). It is interesting to note that, whereas increasing thickness improved the stall of the plain", "timestamp": "2026-07-22T04:57:23.394578+00:00"}
{"citation_id": "19930082487", "source_url": "https://ntrs.nasa.gov/api/citations/19930082487/downloads/19930082487.pdf", "page_number": 33, "total_pages": 33, "image_filename": "19930082487_p33.jpg", "text": "NACA TN No. 1813\n31\n\n<!-- Image (199, 229, 776, 703) -->\n\nFigure 11.- Comparison of calculated and experimental values of local Mach number at airfoil crest for several NACA airfoil sections.", "timestamp": "2026-07-22T04:57:23.888916+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 18, "total_pages": 58, "image_filename": "19930082617_p18.jpg", "text": "NACA TN 1962\n17\n\nStringers\no 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n6 216.0 X $10^3$\n\n3.86\n\n[Figure: Diagram showing Band H and section A-A]\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n-12 -8 -4 0 4 8 12 16 20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6\n\n[Figure: NACA logo]\n\nFigure 6.- Strain diagram of cylinder 73. Band H.", "timestamp": "2026-07-22T04:57:25.084484+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 7, "total_pages": 46, "image_filename": "19930085519_p7.jpg", "text": "6\nNACA RM No. L8K19\n\nThe 20-percent-chord by 49-percent-semispan plain aileron investigated was formed by cutting the flap on a line parallel to the wing plane of symmetry. The plain aileron was sealed and was held at the various deflections (ranging from 20° to -20°) by steel straps on both the wing upper and lower surfaces.\n\nTransition was not fixed for any of the tests.\n\nTESTS\n\nThe slotted-flap, plug-aileron, and plain-aileron tests were performed at an average dynamic pressure of approximately 20.5 pounds per square foot, which corresponds to a Mach number of about 0.12, and a Reynolds number of about 2,400,000. The Zap flap test and the plain-wing test performed in conjunction with the Zap flap test were performed at an average dynamic pressure of about 9.1 pounds per square foot, which corresponds to a Mach number of about 0.07 and a Reynolds number of about 1,600,000. Both Reynolds numbers are based on the wing mean aerodynamic chord of 2.89 feet.\n\nThe tests, in general, were run through a range of angle of attack of -10° to 26°.\n\nRESULTS AND DISCUSSION\n\nThe lift, drag, pitching-moment, and calculated trimmed-gliding characteristics of the 42° sweptback semispan-wing model are presented in figures 10 to 16. The rolling-moment and yawing-moment characteristics of the various plug-aileron and plain-aileron configurations, both flap-neutral and flap-deflected, are presented in figures 17 to 22.\n\nWing Aerodynamic Characteristics\n\nFlap retracted.- The aerodynamic characteristics of the wing with various plug-slot configurations are shown in figure 10. It may be seen from figure 10 that at an angle of attack of about 16°, regardless of the plug-gap or lower-lip configuration, the slope of the pitching-moment-coefficient curve becomes markedly unstable and the drag starts to increase rapidly. A visual study of the behavior of tufts on the upper surface of the wing showed that a sudden stalling of approximately the outboard 40 percent of the wing occurred at this angle of attack. This abrupt stall may be a condition encountered only at the low Reynolds number at which the tests were performed. The results of previous unpublished tests in the Langley 19-foot pressure tunnel of a complete wing (with individual panels having the same geometric characteristics as the wing reported herein)", "timestamp": "2026-07-22T04:57:27.905619+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 10, "total_pages": 28, "image_filename": "19930085471_p10.jpg", "text": "8\nUNCLASSIFIED\nRECONFIDENTIAL\nNACA RM No. L8J11\n\n4. More systematic experimental examination of the effects of the\nindividual parameters is desirable.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Field, Va.\n\nREFERENCES\n\n1. Garrick, I. E., and Rubinow, S. I.: Flutter and Oscillating Air-Force\nCalculations for an Airfoil in a Two-Dimensional Supersonic Flow.\nNACA Rep. No. 846, 1946.\n\n2. Theodorsen, Theodore, and Garrick, I. E.: Mechanism of Flutter -\nA Theoretical and Experimental Investigation of the Flutter\nProblem. NACA Rep. No. 685, 1940.\n\n3. Smilg, Benjamin: An Engineering Evaluation of Flutter and Other\nAero-Elastic Problems at Transonic and Supersonic Speeds. ATI\nTech. Data Digest, vol. 13, no. 14, July 15, 1948, pp. 9-18.\n\n4. Barmby, J. G., and Clevenson, S. A.: Initial Test in the Transonic\nRange of Four Flutter Airfoils Attached to a Freely Falling Body.\nNACA RM No. L7B27, 1947.\n\n5. Barmby, J. G., and Teitelbaum, J. M.: Initial Flight Tests of the\nNACA FR-2, a High-Velocity Rocket-Propelled Vehicle for Transonic\nFlutter Research. NACA RM No. L7J20, 1948.\n\n6. Clevenson, S. A., and Lauten, William T., Jr.: Flutter Investigation\nin the Transonic Range of Six Airfoils Attached to Three Freely\nFalling Bodies. NACA RM No. L7K17, 1948.\n\nUNCLASSIFIED\nCONFIDENTIAL\nRESTRICTED", "timestamp": "2026-07-22T04:57:30.087953+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 18, "total_pages": 30, "image_filename": "19930082585_p18.jpg", "text": "NACA TN 1907\n17\n\nInfluence of Rate of Pitch Reduction and Blade\nMoment of Inertia on $V(t)$ and $\\Omega(t)$\n\nThe influence of rate of pitch change and blade moment of inertia on the variation of flapping angle with time is shown in figure 5; the effect of blade moment of inertia on $V(t)$ and $\\Omega(t)$ is shown in figure 6. It may be considered that the average time required to effect the transition for the cases considered is about 6 seconds. With regard to avoiding the instability caused by blade stalling at high values of $\\lambda$ (see the study of stability of autorotation, reference 1), it is desirable to keep $\\lambda$ to a minimum throughout the maneuver.\n\nIn order to evaluate the effects of rate of pitch reduction and moment of inertia in this regard, figures 7 and 8 are presented. The solid lines represent boundaries determined from reference 1, figure 3. For any value of $\\theta$, the point on the boundary was determined by the value of $\\lambda$ at which the difference in $\\frac{2C_Q}{\\sigma}$ due to neglecting blade stalling was 0.0015. Thus the boundaries shown indicate, for any value of $\\theta$, an arbitrary limit for $\\lambda$, above which blade stalling should be accounted for in equations (12a) and (13a). The broken curves are the loci of combinations of $\\lambda$ and $\\theta$ computed during the maneuvers considered in the study of the sample design. As can be seen from the figures, all the cases computed lie within the range of $\\lambda$ and $\\theta$ where blade stalling can be neglected, as has been done in the ANALYSIS.\n\nFrom figure 7, the advantage of a rapid pitch reduction is apparent, for keeping $\\lambda$ to a minimum throughout the maneuver, and hence in avoiding the instability of excessive blade stalling. From figure 8, a similar advantage for large blade moment of inertia is to be noted.\n\nFor larger initial blade angles in hovering than considered here, the maneuver can only be investigated by accounting for blade stalling, and then only with grave reservations because of the assumption of constant induced velocity, which assumption was shown in reference 1 to be severe in cases where blade stalling is important. It can, however, be anticipated that, in this case, the importance of rapid rate of pitch reduction and large blade moment of inertia would be greatly magnified.\n\nDescending Velocity against Altitude Lost\n\nFigures 9 and 10 show descending velocity against altitude lost during the maneuvers investigated. Altitude lost is obtained by graphical integration of the curves of $V(t)$, figures 5 and 6. The", "timestamp": "2026-07-22T04:57:30.708428+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 32, "total_pages": 41, "image_filename": "19930082476_p32.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:57:31.892369+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 12, "total_pages": 14, "image_filename": "19930082712_p12.jpg", "text": "```markdown\n10\n\n1.6\n1.2\n.8\n.4\n0\n-.4\n-.8\nSection lift coefficient, $c_l$\n\n.2\n0\n-.1\n-.2\nMoment coefficients, $c_{m_{c/4}}$\n\n-8 0 8 16 0 24 0 0 0 0\nSection angle of attack, $\\alpha_s$, deg\n\nR\n$\\circ$ 1.8 x $10^6$\n$\\square$ 2.6\n$\\triangle$ 3.0\n$\\diamond$ 6.0\n$\\nabla$ 9.0\n$\\triangleright$ 11.0\nFlagged symbols denote\nstandard roughness\n\nNACA\n\n(a) Section lift and pitching-moment characteristics.\n\nFigure 1.- Aerodynamic characteristics of the NACA 8-H-12 airfoil section, 24-inch chord.\n\nNACA TN 1998\n```", "timestamp": "2026-07-22T04:57:34.867275+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 20, "total_pages": 49, "image_filename": "19930082498_p20.jpg", "text": "```markdown\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, f (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | | Other sounds | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | Over-all | 0.5f | 1.0f | 1.5f | 2.0f | 2.5f | 3.0f | 3.5f | 4.0f | 4.5f | 5.0f | 5.5f | 6.0f | 6.5f | 7.0f | cps | db | cps | db | |\n| (a)<br>12<br>18<br>18<br>2 3/4<br>25<br>25<br>Muffler attached to 106-inch exhaust pipe and terminated with upturned elbow (see fig. 1) | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 76.0<br>83.0<br>88.5 | 40<br>50<br>--- | 76<br>83<br>--- | 58<br>50<br>--- | 58<br>50<br>--- | 46<br>(b)<br>--- | 50<br>58<br>--- | 47<br>50<br>--- | 45<br>50<br>--- | 45<br>50<br>--- | 50<br>(b)<br>--- | 45<br>(b)<br>--- | 40<br>(b)<br>--- | 40<br>(b)<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | Medium |\n| 2 3/4<br>18<br>18<br>24<br>25<br>25<br>16<br>Baffle welded at 4 spots on circumference | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 91.0<br>88.5<br>91.0 | (b)<br>45<br>(b) | 45<br>66<br>88 | 62<br>66<br>--- | 68<br>55<br>70 | 58<br>60<br>72 | 58<br>81<br>86 | 45<br>63<br>84 | 53<br>63<br>65 | 53<br>70<br>84 | 50<br>67<br>60 | 45<br>80<br>(b) | 45<br>55<br>(b) | (b)<br>(b)<br>(b) | (b)<br>(b)<br>(b) | ---<br>---<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | ---<br>---<br>--- | Medium |\n| 45 Baffle seam welded; otherwise same as 44 (see fig. 2(e)) | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 84.5<br>83.0<br>87.0 | (b)<br>50<br>58 | 73<br>65<br>84 | 50<br>50<br>58 | 51<br>50<br>60 | (b)<br>58<br>65 | 50<br>81<br>84 | 58<br>66<br>72 | 50<br>63<br>63 | 51<br>68<br>66 | 53<br>68<br>66 | 50<br>(b)<br>(b) | 50<br>(b)<br>(b) | (b)<br>(b)<br>(b) | (b)<br>(b)<br>(b) | ---<br>---<br>--- | ---<br>---<br>--- | 450<br>---<br>--- | 82<br>---<br>--- | 70<br>---<br>--- | Medium |\n| 12<br>14<br>14<br>2 3/4<br>25<br>25 | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 77.0<br>81.5<br>86.0 | 35<br>(b)<br>54 | 68<br>67<br>84 | 65<br>58<br>58 | 48<br>50<br>73 | 55<br>54<br>63 | 57<br>59<br>73 | 50<br>58<br>71 | 56<br>69<br>61 | 64<br>59<br>67 | 70<br>53<br>67 | 65<br>48<br>69 | 59<br>48<br>(b) | 53<br>(b)<br>(b) | 58<br>51<br>730 | 63<br>400<br>77 | 750<br>754<br>--- | 55<br>115<br>380 | 55<br>60<br>89 | Low |\n| 12<br>1<br>1<br>2 3/4<br>25<br>25 | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 81.5<br>82.5<br>93.5 | (b)<br>45<br>55 | 71<br>59<br>90 | 45<br>62<br>60 | 63<br>66<br>79 | 45<br>48<br>68 | 45<br>66<br>72 | 53<br>68<br>65 | 62<br>55<br>69 | 59<br>53<br>75 | 53<br>55<br>72 | 50<br>53<br>79 | 45<br>60<br>65 | 58<br>60<br>(b) | 51<br>58<br>375 | 360<br>260<br>61 | 650<br>650<br>730 | 55<br>60<br>390 | 55<br>60<br>71 | Low |\n| 12<br>1<br>2 3/4<br>20<br>24<br>(see fig. 1) | 1650<br>2000<br>2750 | 82.5<br>100.0<br>139.5 | 84.0<br>86.0<br>101.5 | (b)<br>45<br>64 | 75<br>78<br>96 | 63<br>62<br>79 | 81<br>84<br>95 | 67<br>72<br>86 | 70<br>78<br>95 | 71<br>69<br>--- | 73<br>71<br>84 | 63<br>72<br>--- | 63<br>62<br>72 | 63<br>63<br>77 | 57<br>62<br>76 | 56<br>58<br>77 | 58<br>58<br>70 | 742<br>390<br>250 | 62<br>73<br>70 | 783<br>820<br>1400 | 65<br>67<br>73 | Low |\n\nIn the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\nThe sound-pressure level was below the range of the analyzer.\n\nNACA\nNACA TN No. 1838\n61\n```", "timestamp": "2026-07-22T04:57:41.522250+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 17, "total_pages": 36, "image_filename": "19930082614_p17.jpg", "text": "NACA TN 1939\n\ndrag is attributable to the change in angle of attack necessary for the lift to be constant. Because the rate of change of drag with angle of attack for lift coefficients near zero is small, the effect of such an angle-of-attack change is not large for the data presented. Apparently, therefore, the large drag values of the spoiler-type brakes result from changes in the flow over the wing.\n\nFrom the discussion of the functions of aerodynamic brakes, it is evident that they should be designed to be effective throughout a wide range of conditions. There should be a smooth increase in drag as the brake is extended, permitting any position between fully open and closed to be selected with a corresponding control over the deceleration. The variations of drag coefficient with percent extension for brakes of several types are shown in figure 7.\n\nThe effect of Mach number upon air-brake drag is dependent upon the particular installation. The variations of incremental drag coefficient with Mach number for two aerodynamic brakes (of type D) on a rectangular wing are shown in figure 8. The drag characteristics are affected to a large extent by the lift on the wing. At an angle of attack of $-1.0^\\circ$, the rate of rise of drag with Mach number became greater as the Mach number increased from 0.3 to 0.775. At an angle of attack of $3.0^\\circ$, the drag increased with Mach number up to a Mach number of 0.7 and then began to decrease. The drag coefficient due to the fuselage side brake, shown in figure 8, increased nearly uniformly with Mach number throughout the range of Mach numbers from 0.3 to 0.875. The drag coefficient due to the fuselage dive-recovery flaps (type P) increased rapidly with Mach number above 0.6, rising to 161 percent of its low-speed value at a Mach number of 0.8.\n\nEXAMPLE CALCULATIONS\n\nExamples are presented to illustrate the procedures for calculating the variations with time of the forward speed and altitude. These calculations are for an airplane with a wing loading of 50 pounds per square foot, initially flying at a true airspeed of 700 feet per second. An initial altitude of 25,000 feet was assumed and some additional calculations were made assuming an initial altitude of 10,000 feet to indicate the effects of this reduction in altitude.\n\nThe longitudinal aerodynamic forces on the airplane can be represented by the coefficient\n\n$$\nC_{D_n} = C_{D_A} + C_{D_T} + F C_L^2 + \\Delta C_D\n$$\n\nwhere the terms are defined as follows:", "timestamp": "2026-07-22T04:57:43.540363+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 5, "total_pages": 20, "image_filename": "19930085536_p5.jpg", "text": "NACA RM No. E8K05\n\nAPPARATUS AND PROCEDURE\n\nThe cone was mounted on a support body in the Lewis 18-by-18-inch supersonic tunnel, as shown in figure 1. The support body was a sweptback strut fastened to the tunnel wall by means of a lock nut. From a previous calibration, the Mach number in the vicinity of the model was 1.89 with a maximum deviation of ±0.5 percent.\n\nA sketch of the model showing the dimensions and the location of the pressure orifices is presented in figure 2. The body was machined of brass and the nose was finished to a sharp point. Orifices of 0.010-inch diameter were drilled normal to the body surface. Pressures were photographically recorded on a multiple-tube manometer board using tetrabromethane as a fluid.\n\nThe model, mounted as shown in figure 1, and the strut were turned together to obtain data for the body in yaw. In order to obtain the desired angle of attack, the model was rotated $90^\\circ$ relative to the strut and the angle was varied by turning the strut. By use of a vernier, the angle could be read to within 2.5 minutes. Pressures were recorded every $0.5^\\circ$ up to $\\pm 16^\\circ$ angle of yaw at an angle of attack of $0^\\circ$ and $\\pm 10^\\circ$ angle of attack at an angle of yaw of $0^\\circ$.\n\nTHEORY\n\nA method of calculating the pressure distribution about a cone of arbitrary cross section by means of a series of line sources is presented in reference 6. The following perturbation velocities result from such sources:", "timestamp": "2026-07-22T04:57:50.014864+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 27, "total_pages": 99, "image_filename": "19930082511_p27.jpg", "text": "NACA TN No. 1826\n25\n\nsatisfies the preceding conditions and has a pole of the first order\nat $z_1$. $Q(z)$ is therefore of the form:\n\n$$Q(z) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) (A + Bz) \\sqrt{\\frac{1 - z^2}{a - z}} \\quad (10)$$\n\nThe real constants A and B are determined by the same condition as\nbefore, which gives\n\n$$\\Gamma' = - \\frac{2(A + Bz_1)}{z_1} \\sqrt{\\frac{1 - z_1^2}{a - z_1}} \\quad (11)$$\n\nTunnel-interference velocity.- The interference velocity $q_1(\\xi, \\xi_1)$\nis the difference between the perturbation velocity and the velocity due\nto a vortex located at the same point in an unbounded medium:\n\n$$q_1(\\xi, \\xi_1) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) (A + Bz) \\sqrt{\\frac{1 - z^2}{a - z}} + \\frac{i\\Gamma'}{2\\pi(\\xi - \\xi_1)} \\quad (12)$$\n\nThis expression does not simplify appreciably if the vortex is located\non the axis, and the interference velocity at a point on the axis due\nto a vortex on the axis is not normal to the axis.\n\nThe interference velocity at the vortex itself is the limit of\nexpression (12) as $z$ approaches $z_1$. Proceeding as in the preceding\ncase gives:\n\n$$q_1(\\xi_1) = i \\left[ \\frac{\\Gamma'}{2(1 - z_1^2)} - \\frac{\\Gamma' a}{4(a - z_1)} + B \\sqrt{\\frac{1 - z_1^2}{a - z_1}} + \\frac{\\Gamma' z_1}{4i y_1} \\right] \\quad (13)$$\n\nIt may be shown with the aid of equation (11) that this equation reduces\nto that for the closed-open tunnel as $a$ goes to infinity.\n\nCASE 3 - CLOSED-OPEN-CLOSED TUNNEL\n\nPerturbation velocity.- The transformation $z = e^{\\pi \\xi}$ transforms the\ntunnel space into the upper half of the $z$-plane with correspondence\nbetween points as indicated in figure 15. The boundary conditions on\nthe complex velocity $Q(z)$ are:", "timestamp": "2026-07-22T04:57:51.437796+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 13, "total_pages": 66, "image_filename": "19930082914_p13.jpg", "text": "12\nNACA TN No. 1857\n\nthe fringes are located between $S_2$ and the telescope and are too narrow. The first step is to determine which of the two beams goes through which path in the interferometer. This is determined by focusing the telescope on the light source and blocking off one of the paths. It is desired to have the beams cross in the test section, midway between $M_2$ and $S_2$.\n\nIf the slope of the beam that passes through the test section is positive with respect to the slope of the other beam, as shown in figure 4, then plate $S_2$ is rotated counterclockwise to bring the fringes back almost to the test section. This adjustment may make the fringes too broad. Then $S_1$ is rotated clockwise to move the fringes the rest of the way to the test section and at the same time to narrow them.\n\nThe next adjustment is to make the two optical-path lengths through the interferometer equal. With the fringes in focus in or near the test section, the slits are opened until the fringes visible in the telescope become faint. By rotating a compensating plate, the fringes are caused to appear to pass vertically through the test section until they disappear. The position of the compensating plate is noted, and the plate is rotated in the opposite direction until the fringes with the greatest contrast have passed through the test section and the fringes again disappear. The compensating plate is then positioned halfway between the two positions where the fringes disappeared. The interferometer is then nearly in adjustment for white-light fringes. Either the monochromator is removed and a source of white light is used, or the white light is placed just ahead of the first splitter plate without disturbing the monochromator. (A convenient point source is the zirconium-arc light.) Then a slight adjustment of a compensating plate will bring the white-light fringes into view. If the disturbance of which an interferogram is to be taken contains a region in which there is a density gradient of considerable magnitude, such as a boundary layer, the fringes will be crowded together in that region, and distinguishing individual fringes may be difficult. It is advisable to place the white-light fringes in such a position that they will move, when the disturbance is produced, into the region of density gradient and thereby provide the fringes of greatest contrast in that region.\n\nThe final adjustment is to remove what might be called \"twist\" from the two beams. For horizontal fringes the two splitter plates have been rotated about horizontal axes and the two images of the light source, as seen in the telescope, lie one above the other. It may be, however, that the two beams do not lie in the same vertical plane. If they do not, sharp fringes can be observed only when the source is a line of zero width (and is also vertical, for horizontal fringes). As it is necessary that the light source have a finite width, in order to give enough light, then the fringes will appear considerably blurred unless the rotation of the plates is corrected in order that the two beams lie in the same vertical plane. To check this, the telescope is removed and the eye placed in the emergent beam some distance away from the interferometer.", "timestamp": "2026-07-22T04:57:57.198696+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 35, "total_pages": 62, "image_filename": "19930082485_p35.jpg", "text": "34\nNACA TN No. 1810\n\n[Figure: Diagram of a curved section showing survey paths, circumferential distance, and radius measurements. Labels include: Outer shroud, Inner shroud, Survey path, Circumferential distance, Radius. Numerical values shown: 0, 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 11.7]\n\nNACA\n\nFigure 4.- Survey planes 1.5 chord lengths upstream of the turbine blades. (All dimensions are given in inches.)", "timestamp": "2026-07-22T04:57:57.284633+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 40, "total_pages": 47, "image_filename": "19930093773_p40.jpg", "text": "```markdown\nNACA RM E9G09\n39\n\nCorrected fuel-air ratio, (f/a)/$\\theta$\n\n| Altitude (ft) | |\n| :--- | :--- |\n| $\\circ$ | 5,000 |\n| $\\diamond$ | 15,000 |\n| $\\diamond$ | 25,000 |\n| $\\triangle$ | 35,000 |\n| $\\triangle$ | 45,000 |\n| $\\nabla$ | 50,000 |\n\n.024\n.020\n.016\n.012\n.008\n.004\n0\n\n3 4 5 6 7 8 9x $10^3$\nCorrected engine speed, N/$\\sqrt{\\theta}$, rpm\n\n[NACA logo]\n\n(e) Fuel-air ratio.\nFigure 6. - Continued. Effect of altitude on variation of corrected engine performance with corrected engine speed at flight Mach number of 0.21.\n```", "timestamp": "2026-07-22T04:58:02.554938+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 1, "total_pages": 46, "image_filename": "19930085542_p1.jpg", "text": "NACA RM No. L8L29\n\nRM No. L8L29\n\n[Figure: NACA logo with wings]\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED STATIC-STABILITY AND ROLLING CHARACTERISTICS OF  \nLOW-ASPECT-RATIO WINGS OF TRIANGULAR AND  \nMODIFIED TRIANGULAR PLAN FORMS\n\nBy  \nByron M. Jaquet and Jack D. Brewer\n\nLangley Aeronautical Laboratory  \nLangley Air Force Base, Va.\n\nNATIONAL ADVISORY COMMITTEE  \nFOR AERONAUTICS  \nWASHINGTON\n\nMarch 29, 1949  \nDeclassified December 14, 1953", "timestamp": "2026-07-22T04:58:09.343223+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 19, "total_pages": 58, "image_filename": "19930082617_p19.jpg", "text": "```markdown\n18\nNACA TN 1962\n\nStringers\nO 1 to 9\nX 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 72.0 X $10^3$\n3 108.0 X $10^3$\n4 144.0 X $10^3$\n5 180.0 X $10^3$\n6 216.0 X $10^3$\n\n3.86\"\nA\nBand N\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n12 8 4 0 4 8 12 16 20 X $10^{-4}$\nStrain\n\n1 2 3 4 5 6\n\nNACA\n\nFigure 7.- Strain diagram of cylinder 73, Band N.\n```", "timestamp": "2026-07-22T04:58:11.647564+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 17, "total_pages": 78, "image_filename": "19930082618_p17.jpg", "text": "NACA TN 1945\n\nNACA 6-series sections, the NACA $6_{42}-415$ and NACA $6_{43}-418$ sections with flaps have stalling characteristics at Reynolds numbers below $6.0 \\times 10^6$ generally less desirable than those of the 9-percent-thick section. Neither variations in thickness form nor in amount of camber seem to have a very important effect upon the stall, although the stall of the 66-series section at the different Reynolds numbers may be somewhat more desirable than the stalls of the other airfoils of 15-percent thickness. The smooth NACA 23012, 23015, and 0012 airfoil sections with split flaps are characterized by quite abrupt stalls at all Reynolds numbers. The NACA 44-series sections, when equipped with split flaps, possess stalling properties which are somewhat similar to those of comparable NACA 6-series sections.\n\nExcept for the NACA 0012 and NACA 230-series sections, the addition of roughness usually improves to some degree the stalling characteristics of the airfoils with flaps, although this is not always true. (See, for example, the data for the NACA $6_{43}-418$ section, fig. 4(b).)\n\nPitching Moment and Aerodynamic Center\n\nThe values of the quarter-chord pitching-moment coefficient corresponding to the design angles of attack show practically no variation with Reynolds number for any of the plain airfoils (figs. 1 to 15, part (c)). Accompanying changes in the Reynolds number, some change in the slope of the pitching-moment curve against angle of attack is noticeable. Consequently, the chordwise position of the aerodynamic center varies somewhat with Reynolds number; however, these variations do not appear to form any consistent trend with the Reynolds number (figs. 1 to 15).\n\nWhen the airfoils are equipped with split flaps there is some variation with Reynolds number of the quarter-chord pitching moment corresponding to zero angle of attack for several of the airfoils. This variation usually consists of a decrease in magnitude of the coefficient with decreasing Reynolds number and is most pronounced for the thicker airfoils with far back position of minimum pressure. There is also some change in the shape of the curve of pitching moment plotted against angle of attack with Reynolds number for several of the airfoils. This change of shape usually consists of a decrease in magnitude of the pitching moment with increasing angle of attack which becomes more pronounced as the Reynolds number is reduced. The magnitude of the effect seems to become more pronounced as the airfoil thickness and camber are increased and as the position of minimum pressure is moved rearward.", "timestamp": "2026-07-22T04:58:18.907685+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 19, "total_pages": 53, "image_filename": "19930082542_p19.jpg", "text": "18\nNACA TN No. 1867\n\nHigher solution temperatures cause lower strengths and extreme brittleness in the rupture test.\n\n4. Incomplete data suggest that further testing would show distinct benefits from aging after hot-cold-work.\n\n5. Properly hot-worked, or hot-worked and then hot-cold-worked, material will have the best all-round properties. Such treatments are not recommended, however, because of the difficulty of controlling hot-working conditions. The original hot-worked material could vary from the fully solution-treated to the severely hot-cold-worked condition depending on the hot-working conditions.\n\nThere are certain limitations to the data. Testing was kept to the bare minimum to show trends. Other characteristics than physical properties at room temperature and rupture properties at $1200^\\circ$ F have not been considered. More work on other heats is needed to verify the reliability of the reported data. More complete design data are needed for the optimum treatments. The optimum treatments for service at higher temperatures than $1200^\\circ$ F are certain to be different from those found in this investigation.\n\nThe trends found for the effects of various treatments in this investigation should apply to other alloys of the same type. Optimum treatment conditions, however, will almost certainly vary for each alloy.\n\nRough general relationships exist between properties and Brinell hardness. By the use of these relationships and the detailed effects of various treatments the order of magnitude of test results from large forged discs can be predicted quite closely. The properties of discs will in general be somewhat lower than those of bar stock.\n\nThis investigation did not include a careful study of the mechanism by which the treatments affect properties. It appears, however, that precipitation has very little influence on room-temperature properties. The highest rupture properties result from precipitation during working or testing at $1200^\\circ$ F. Aging at higher temperatures adversely affects the precipitation reaction. High strength at room temperature and high rupture strength for short time periods after hot-cold-working at temperatures up to $1200^\\circ$ F are apparently due to strain hardening. Strain hardening probably results in very effective precipitation during rupture testing. The evidence regarding precipitation resulting from or during working at higher temperatures than $1200^\\circ$ F is not clear but apparently working temperatures above $1400^\\circ$ F are unfavorable to $1200^\\circ$ F strengths.\n\nUniversity of Michigan\nAnn Arbor, Mich., September 5, 1947.", "timestamp": "2026-07-22T04:58:19.117996+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 19, "total_pages": 30, "image_filename": "19930082585_p19.jpg", "text": "18\nNACA TN 1907\n\naverage drop required to effect the transition, for the cases considered,\nis seen to be about 120 feet.\n\nThe effect of rate of pitch change on descending velocity at a\ngiven amount of altitude lost, from figure 9, is apparently small. This\nconclusion is obviously limited to cases where the hovering incidence is\nwithin the range for steady autorotation. If $\\theta_0$ were greater than the\nmaximum for steady autorotation (reference 1, fig. 3), then obviously\nthe pitch would have to be reduced rapidly to avoid the instability\ncaused by excessive blade stalling.\n\nFrom figure 10, the advantage of a large blade moment of inertia, in\npreserving a minimum descending velocity for a given altitude lost, is\nreadily apparent.\n\nCONCLUSIONS\n\nAnalytical methods are given for treating the transient motion of\na helicopter between the hovering state and the steady autorotative\nstate of vertical descent. From the standpoint of avoiding blade\nstalling in the transition, it is desirable to reduce pitch rapidly and\nto have a large blade moment of inertia.\n\nThe effects of hinging the blades, for a given net rate of pitch\nreduction after power failure, appear to be negligible.\n\nThe average time to effect the transition from hovering to steady\nautorotation is about 6 seconds, for the cases investigated. The\ncorresponding average altitude lost in the transition is about 120 feet.\n\nPrinceton University\nPrinceton, N. J., May 12, 1948", "timestamp": "2026-07-22T04:58:23.483674+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 18, "total_pages": 36, "image_filename": "19930082614_p18.jpg", "text": "16\nNACA TN 1939\n\n$C_{D_A}$ drag coefficient of the airplane without air brakes, excluding the induced drag\n\n$C_{D_T}$ force coefficient, either drag or thrust, due to the propulsion unit\n\nF induced drag factor\n\n$\\Delta C_D$ increment in drag coefficient resulting from the extension of aerodynamic brakes\n\nLevel Flight\n\nThe speed variation for level flight is given by equation (2) in which it is necessary that $C_{D_n}$ be constant.\n\nThe following coefficients have been assumed:\n\n$$\n\\begin{aligned}\nC_{D_A} + C_{D_T} &= 0.013 \\\\\nF &= 0.060 \\\\\nFC_L^2 &= 0.001 \\\\\n\\Delta C_D &= 0.100 \\\\\nC_{D_n} &= 0.114\n\\end{aligned}\n$$\n\nFor an altitude of 25,000 feet,\n\n$$\n\\begin{aligned}\nK &= C_{D_n} \\frac{\\rho}{2} \\frac{g}{W/S} \\\\\n&= (0.114) \\frac{0.001066}{2} \\left( \\frac{32.2}{50} \\right) = 0.0391 \\times 10^{-3} \\text{ per foot} \\\\\nV &= \\frac{1}{Kt + (1/V_o)} = \\frac{25,600}{t + 36.6}\n\\end{aligned}\n$$\n\nFor an altitude of 10,000 feet", "timestamp": "2026-07-22T04:58:23.688904+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 8, "total_pages": 36, "image_filename": "19930085487_p8.jpg", "text": "6\nNACA RM No. E8J22\n\nTABLE I\nFUNDAMENTAL NATURAL FREQUENCIES OF BLADES ON COMPRESSOR ROTOR\n\n<!-- Table (158, 125, 890, 966) -->\n\\begin{tabular}{|l|*{10}{c|}}\n\\hline\n\\multicolumn{1}{|c|}{\\multirow{3}{*}{Blade}} & \\multicolumn{10}{c|}{Natural frequencies} \\\\\n\\cline{2-11}\n\\multicolumn{1}{|c|}{} & \\multicolumn{10}{c|}{(cps)} \\\\\n\\cline{2-11}\n\\multicolumn{1}{|c|}{} & \\multicolumn{10}{c|}{Stages} \\\\\n\\hline\n\\multicolumn{1}{|c|}{} & 1 & 2 & 3 & 4 & 5 & 6 & 7 & 8 & 9 & 10 \\\\\n\\hline\n1 & 386 & 492 & 556 & 580 & 726 & 758 & 920 & 1076 & 1195 & 1400 \\\\\n2 & 378 & 484 & 556 & 620 & 712 & 774 & 884 & 1182 & 1276 & 1350 \\\\\n3 & 416 & 476 & 570 & 616 & 702 & 802 & 936 & 954 & 1030 & 1340 \\\\\n4 & 382 & 474 & 544 & 568 & 690 & 802 & 936 & 1000 & 1276 & 1470 \\\\\n5 & 400 & 490 & 478 & 604 & 734 & 804 & 910 & 1056 & 1204 & 1480 \\\\\n6 & 396 & 488 & 494 & 586 & 740 & 796 & 896 & 1094 & 984 & 1430 \\\\\n7 & 394 & 486 & 548 & 566 & 724 & 808 & 974 & 1120 & 1270 & 1470 \\\\\n8 & 374 & 502 & 514 & 656 & 750 & 764 & 890 & 1040 & 1202 & 1400 \\\\\n9 & 394 & 498 & 480 & 652 & 732 & 814 & 886 & 980 & 1116 & 1220 \\\\\n10 & 394 & 498 & 506 & 574 & 654 & 770 & 900 & 1160 & 1136 & 1390 \\\\\n11 & 400 & 486 & 512 & 588 & 696 & 814 & 930 & 1116 & 1130 & 1370 \\\\\n12 & 398 & 498 & 522 & 598 & 716 & 796 & 922 & 1100 & 1116 & 1470 \\\\\n13 & 396 & 494 & 494 & 574 & 700 & 802 & 880 & 1016 & 1040 & 1330 \\\\\n14 & 390 & 474 & 522 & 594 & 694 & 792 & 888 & 1090 & 976 & 1320 \\\\\n15 & 372 & 514 & 512 & 578 & 688 & 754 & 890 & 1080 & 1130 & 1460 \\\\\n16 & 412 & 476 & 486 & 578 & 690 & 826 & 903 & 974 & 1080 & 1400 \\\\\n17 & 362 & 466 & 522 & 566 & 680 & 774 & 844 & 1070 & 1120 & 1440 \\\\\n18 & 376 & 472 & 494 & 594 & 724 & 808 & 908 & 1060 & 1154 & 1220 \\\\\n19 & 396 & 490 & 512 & 572 & 742 & 810 & 940 & 1026 & 1300 & Loose \\\\\n20 & 392 & 482 & 506 & 584 & 720 & 806 & 910 & 1020 & 1200 & 1200 \\\\\n21 & 390 & 486 & 524 & 568 & 714 & 780 & 856 & 1020 & 1240 & Loose \\\\\n22 & 390 & 488 & 466 & 562 & 724 & 784 & 882 & 1050 & 1100 & Loose \\\\\n23 & 426 & 454 & 500 & 630 & 718 & 800 & 880 & 1020 & 1200 & Loose \\\\\n24 & & & 502 & 568 & 714 & 718 & 902 & 1100 & 1160 & 1380 \\\\\n25 & & & 526 & 574 & 736 & 824 & 918 & 1080 & 1030 & Loose \\\\\n26 & & & 524 & 588 & 764 & 786 & 886 & 1108 & 1132 & Loose \\\\\n27 & & & 484 & 610 & 650 & 804 & 922 & 1080 & 1242 & 1210 \\\\\n28 & & & 498 & 610 & 688 & 774 & 930 & 1088 & 1110 & 1350 \\\\\n29 & & & 516 & 566 & 712 & 796 & 900 & 1090 & 1180 & 1350 \\\\\n30 & & & 506 & 584 & 682 & 788 & 844 & 1092 & 1100 & 1440 \\\\\n31 & & & 504 & 596 & 696 & 754 & 888 & 1090 & 1220 & 1410 \\\\\n32 & & & 514 & 568 & 686 & 778 & 892 & 1146 & 1200 & 1360 \\\\\n33 & & & & & 748 & 814 & 958 & 1080 & 1040 & 1350 \\\\\n34 & & & & & 696 & 818 & 920 & 1100 & 1244 & 1430 \\\\\n35 & & & & & & 806 & 906 & 1040 & 1102 & 1480 \\\\\n36 & & & & & & 822 & 900 & 1070 & 1204 & 1330 \\\\\n37 & & & & & & 810 & 914 & 1002 & 1300 & 1360 \\\\\n38 & & & & & & 794 & 900 & 1008 & 1080 & 1350 \\\\\n39 & & & & & & 762 & 866 & 1038 & 1280 & 1330 \\\\\n40 & & & & & & 770 & 918 & 1096 & 1070 & 1330 \\\\\n41 & & & & & & 804 & 926 & 1118 & 1360 & 1390 \\\\\n42 & & & & & & 820 & 910 & 968 & 1090 & 1480 \\\\\n43 & & & & & & 790 & 898 & 1090 & 1222 & 1330 \\\\\n44 & & & & & & & 820 & 1114 & 1092 & 1430 \\\\\n45 & & & & & & & 916 & 1018 & 1030 & 1430 \\\\\n46 & & & & & & & 1028 & 1088 & 1120 & 1420 \\\\\n47 & & & & & & & 1134 & 1070 & 1080 & 1380 \\\\\n48 & & & & & & & & 1120 & 1120 & 1440 \\\\\n49 & & & & & & & & 1038 & 1088 & 1440 \\\\\n50 & & & & & & & & 1090 & 1090 & 1410 \\\\\n51 & & & & & & & & 1030 & 1030 & 1430 \\\\\n52 & & & & & & & & 1140 & 1140 & 1430 \\\\\n53 & & & & & & & & 1070 & 1070 & 1470 \\\\\n54 & & & & & & & & 1120 & 1120 & 1450 \\\\\n55 & & & & & & & & 1028 & 1028 & 1450 \\\\\n56 & & & & & & & & 1170 & 1170 & 1280 \\\\\n57 & & & & & & & & 1140 & 1140 & 1400 \\\\\n58 & & & & & & & & 1240 & 1240 & 1290 \\\\\n59 & & & & & & & & 1050 & 1050 & 1400 \\\\\n60 & & & & & & & & 1210 & 1210 & 1320 \\\\\n61 & & & & & & & & 1100 & 1100 & 1260 \\\\\n62 & & & & & & & & 1090 & 1090 & 1280 \\\\\n63 & & & & & & & & 1140 & 1140 & 1250 \\\\\n64 & & & & & & & & 1180 & 1180 & 1400 \\\\\n65 & & & & & & & & 1250 & 1250 & 1380 \\\\\n66 & & & & & & & & 1120 & 1120 & 1310 \\\\\n67 & & & & & & & & 1150 & 1150 & 1280 \\\\\n68 & & & & & & & & 1120 & 1120 & 1460 \\\\\n69 & & & & & & & & 1140 & 1140 & 1380 \\\\\n70 & & & & & & & & 1040 & 1040 & 1360 \\\\\n71 & & & & & & & & 1130 & 1130 & 1480 \\\\\n72 & & & & & & & & Loose & Loose & 1350 \\\\\n73 & & & & & & & & Loose & Loose & 1310 \\\\\n74 & & & & & & & & Loose & Loose & 1420 \\\\\n75 & & & & & & & & 1088 & 1088 & 1410 \\\\\n76 & & & & & & & & 1130 & 1130 & 1450 \\\\\n77 & & & & & & & & 1080 & 1080 & 1390 \\\\\n78 & & & & & & & & 1120 & 1120 & 1340 \\\\\n\\hline\nHighest & 416 & 514 & 570 & 650 & 764 & 826 & 990 & 1146 & 1380 & 1510 \\\\\nLowest & 352 & 422 & 466 & 556 & 650 & 718 & 856 & 964 & 976 & 1200 \\\\\nAverage & 388 & 481 & 511 & 586 & 712 & 780 & 913 & 1065 & 1146 & 1384 \\\\\n\\hline\n\\end{tabular}\n\nNACA", "timestamp": "2026-07-22T04:58:29.310500+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 8, "total_pages": 46, "image_filename": "19930085519_p8.jpg", "text": "NACA RM No. L5K19\n\nthrough a large range of Reynolds number indicated that at the higher Reynolds number, the break in the pitching-moment curve would be delayed to a higher angle of attack and the abrupt nature of the stall would be somewhat relieved.\n\nFigure 10 shows little variation of the wing aerodynamic characteristics with variation in plug-slot configuration other than that a higher drag was obtained with the plug-slot configuration having the faired plug-slot lower lip and with both the upper and lower plug-slot gaps open than with any of the other plug-slot configurations investigated.\n\nFlap deflected.- A series of flap-nose positions was investigated with the full-span slotted flap at deflections of $30^\\circ$, $40^\\circ$, and $50^\\circ$. The results obtained for the most satisfactory flap positions at each particular deflection are presented in figure 11 and show that all three flap configurations gave about the same value of maximum lift. A flap deflection of $30^\\circ$ in the position indicated gave the highest L/D ratio throughout the lift range and is therefore considered to be the optimum flap deflection.\n\nThe inboard 51-percent-span slotted flap at $50^\\circ$ deflection is considerably more than half as effective in producing lift as the full-span slotted flap at the same deflection and position. (See fig. 12.) The proportionately higher lifting effectiveness of the inboard partial-span flap as compared with the full-span flap has been noted previously in references 4, 5, and 6 for unswept wings. The pitching-moment coefficients produced by the full-span slotted flap is about three times as great as that produced by the partial-span slotted flap. This effect is as would be predicted on the basis of the analyses of references 7 and 8 which show that the center of load of the wing with the full-span flap is considerably farther behind the aerodynamic center of the wing than is the center of load of the wing with the partial-span inboard flap.\n\nFigure 13 presents a comparison of the aerodynamic characteristics in pitch of the plain wing and the wing with the half-span Zap flap deflected $60^\\circ$. (These data were obtained at a Reynolds number of 1,600,000.) A comparison of the plain-wing data of figure 13 with the plain-wing data of figure 12 (which was obtained at a Reynolds number of 2,400,000) shows that the maximum value of $C_L$ and the slope of the curve of $C_L$ against $\\alpha$ at the high lifts obtained at the lower Reynolds number are somewhat greater than the values obtained at the higher Reynolds number. In addition, a comparison of the Zap flap data of figure 13 with slotted-flap data of figure 12 shows that the Zap flap produced almost as high a maximum value of $C_L$ as the full-span slotted flap but produced about the same increment of $C_L$ as the partial-span slotted flap at lift coefficients below maximum lift. The differences between the plain-wing data at the two Reynolds numbers are probably caused by changes in the physical", "timestamp": "2026-07-22T04:58:33.182659+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 36, "total_pages": 62, "image_filename": "19930082485_p36.jpg", "text": "1028\n\nNACA TN NO. 1810\n\nStagnation pressure\n$P_t$, lb/sq ft\n\nRatio of stagnation-\nto-ambient-static\npressure, $P_t/P_{s,0}$\n\n$\\circ$ 1.124\n$\\square$ 1.270\n$\\diamond$ 1.523\n$\\triangle$ 1.682\n\nStatic pressure\n$P_s$, lb/sq ft\n\n(a) Circumferential distance, in.\n\nStagnation pressure\n$P_t$, lb/sq ft\n\nInner shroud\nOuter shroud\n\nStatic pressure\n$P_s$, lb/sq ft\n\nInner shroud\nOuter shroud\n\n(b) Radius, in.\n\nFigure 5. - Variation of stagnation and static pressures in the cascade tunnel at 1.5 chord lengths upstream of the turbine blades.\n\n35", "timestamp": "2026-07-22T04:58:33.584352+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 28, "total_pages": 99, "image_filename": "19930082511_p28.jpg", "text": "26\nNACA TN No. 1826\n\n(1) On the real axis, I.P.Q(z) = 0 for $|z| < 1$ and for $|z| > a$\n(2) On the real axis, R.P.Q(z) = 0 for $1 < |z| < a$\n(3) For $z = \\pm 1$, $Q(z) = 0$\n(4) Q(z) is finite everywhere in the upper half-plane except at $z = \\pm a$ and at $z = z_1$\n(5) $Q(0) = Q(\\infty)$, (this equation corresponding to the condition noted in part I that the perturbation velocities in the upstream and downstream closed regions be the same)\n\nThe function G(z) given by equation (2) is again used as a factor of Q(z). The functions $\\sqrt{\\frac{1 - z^2}{a^2 - z^2}}$, $z\\sqrt{\\frac{1 - z^2}{a^2 - z^2}}$, and $z^2\\sqrt{\\frac{1 - z^2}{a^2 - z^2}}$ satisfy conditions (1), (2), and (3), and are of orders 1, z, and $z^2$ at infinity, respectively. By the same reasoning as before,\n\n$$Q(z) = i\\left(\\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1}\\right)(A + Bz + Cz^2)\\sqrt{\\frac{1 - z^2}{a^2 - z^2}} \\quad (14)$$\n\nThe condition that the pole of the first order at $z_1$ represents a vortex of strength $\\Gamma'$ is:\n\n$$\\Gamma' = -\\frac{2(A + Bz_1 + Cz_1^2)}{z_1}\\sqrt{\\frac{1 - z_1^2}{a^2 - z_1^2}} \\quad (15)$$\n\nCondition (5) is satisfied by equating the two forms of equation (14) for z equal to zero and equal to infinity. Thus\n\n$$Q(0) = i\\left(-\\frac{1}{z_1} + \\frac{1}{\\bar{z}_1}\\right)\\frac{A}{a}$$\n\n$$= \\frac{-2y_1}{|z_1|^2}\\frac{A}{a}$$\n\n$$\\lim_{z \\to \\infty} Q(z) = \\lim_{z \\to \\infty} \\frac{-2y_1}{(z - z_1)(z - \\bar{z}_1)}(A + Bz + Cz^2)\\sqrt{\\frac{1 - z^2}{a^2 - z^2}}$$\n\n$$= -2y_1C$$", "timestamp": "2026-07-22T04:58:36.617160+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 14, "total_pages": 66, "image_filename": "19930082914_p14.jpg", "text": "NACA TN No. 1857\n13\n\nThe fringes may then appear to be cocked at an angle to the horizontal. If so, then as one walks toward the interferometer the fringes will slowly rotate back to the horizontal position which they had when viewed with the telescope focused on the test section. By alternately rotating the two plates about their non-horizontal axes, the two beams can be swung until they both lie in the same vertical plane and the fringes appear horizontal when viewed from both a close and a far position.\n\nWhen all these adjustments have been made, the following will be true:\n\n(a) The fringes are centered in the test section.\n(b) The fringes have the desired orientation.\n(c) The fringes have the desired spacing.\n(d) The white-light fringes are in the correct location in a vertical cross section through the test section.\n(e) The two beams lie in a single plane.\n\nEvaluation of Density Fields\n\nThe theory of the evaluation of interferograms of one- and two-dimensional flow fields given here is no different from that given elsewhere. (See reference 11.) The theory is repeated here for the sake of convenience. The method of measuring fringe shifts is believed to be somewhat different.\n\nThe theory of the formation of the fringes has been discussed in a previous section. The production of fringe shifts is illustrated in figure 9. Consider that when no disturbance is present both beams travel in air of density $\\rho$ and refractive index $n$. The value of the refractive index is a function of the density of the air, for a given wave length of light, according to the Lorentz-Lorenz relation\n\n$$\n\\frac{n^2 - 1}{n^2 + 2} \\propto \\rho\n$$\n\nor\n\n$$\n\\frac{n + 1}{n^2 + 2} (n - 1) \\propto \\rho\n$$", "timestamp": "2026-07-22T04:58:36.924326+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 41, "total_pages": 47, "image_filename": "19930093773_p41.jpg", "text": "40\nNACA RM E9G09\n\nCorrected exhaust-gas total temperature, $T_7/\\theta$, $^\\circ$R\n\n| Altitude (ft) | |\n| :--- | :--- |\n| $\\circ$ | 5,000 |\n| $\\square$ | 15,000 |\n| $\\diamond$ | 25,000 |\n| $\\nabla$ | 35,000 |\n| $\\triangle$ | 45,000 |\n| $\\blacktriangle$ | 50,000 |\n\n| | | | | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| 2200 | | | | | | | | |\n| 2000 | | | | | | | | |\n| 1800 | | | | | | | | |\n| 1600 | | | | | | | | |\n| 1400 | | | | | | | | |\n| 1200 | | | | | | | | |\n| 1000 | | | | | | | | |\n| | 3 | 4 | 5 | 6 | 7 | 8 | 9x$10^5$ | |\n\nCorrected engine speed, $N/\\sqrt{\\theta}$, rpm\n\n(f) Exhaust-gas total temperature\n\n[NACA logo]\n\nFigure 6. - Concluded. Effect of altitude on variation of corrected engine performance with corrected engine speed at flight Mach number of 0.21.", "timestamp": "2026-07-22T04:58:40.450810+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 2, "total_pages": 46, "image_filename": "19930085542_p2.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-22T04:58:41.328650+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 25, "total_pages": 37, "image_filename": "19930082450_p25.jpg", "text": "24\nNACA TN No. 1778\n\n$$\n\\frac{P_l}{L\\sqrt{R}}, \\text{ksi}\n$$\n\n| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | 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| | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | | |", "timestamp": "2026-07-22T04:58:45.781815+00:00"}
{"citation_id": "19930085536", "source_url": "https://ntrs.nasa.gov/api/citations/19930085536/downloads/19930085536.pdf", "page_number": 6, "total_pages": 20, "image_filename": "19930085536_p6.jpg", "text": "$$\nU _ { x } = - \\textrm { K } \\left( \\frac { 1 } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } \\cos ^ { - 1 } \\left\\{ \\frac { 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) } { \\sqrt { \\left[ 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } } \\right\\} \\right.\n$$\n\n$$\n\\left. - \\frac { \\left( \\mathrm { m } \\beta \\right) ^ { 2 } - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) } { \\left[ 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } \\sqrt { \\frac { 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } } \\right)\n$$\n\n$$\nU _ { x } = \\textrm { K } \\beta \\left( \\frac { \\mathrm { m } \\beta \\sin \\left( \\theta - \\delta \\right) } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } \\cosh ^ { - 1 } \\left\\{ \\frac { 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) } { \\sqrt { \\left[ 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } } \\right\\} \\right.\n$$\n\n$$\n\\left. + \\frac { \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] - \\left[ 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] \\mathrm { m } \\beta \\sin \\left( \\theta - \\delta \\right) } { \\left[ 1 - \\textrm { m } \\frac { \\textrm { r } } { \\textrm { x } } \\beta ^ { 2 } \\sin \\left( \\theta - \\delta \\right) \\right] ^ { 2 } - \\left[ 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } \\right] \\left[ 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } \\right] } \\sqrt { \\frac { 1 - \\left( \\frac { \\textrm { r } \\beta } { \\textrm { x } } \\right) ^ { 2 } } { 1 - \\left( \\mathrm { m } \\beta \\right) ^ { 2 } } } \\right)\n$$", "timestamp": "2026-07-22T04:58:48.695944+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 20, "total_pages": 30, "image_filename": "19930082585_p20.jpg", "text": "NACA TN 1907\n19\n\nREFERENCE\n\n1. Nikolsky, A. A., and Seckel, Edward: An Analytical Study of the\nSteady Vertical Descent in Autorotation of Single-Rotor\nHelicopters. NACA TN 1906, 1949.", "timestamp": "2026-07-22T04:58:51.997043+00:00"}
{"citation_id": "19930082542", "source_url": "https://ntrs.nasa.gov/api/citations/19930082542/downloads/19930082542.pdf", "page_number": 20, "total_pages": 53, "image_filename": "19930082542_p20.jpg", "text": "NACA TN No. 1867\n19\n\nREFERENCES\n\n1. White, A. E., Freeman, J. W., and Rote, F. B.: Physical Data on Certain Alloys for High Temperature Applications. NACA ACR No. 3D28, 1943.\n\n2. Freeman, J. W., Rote, F. B., and White, A. E.: High Temperature Characteristics of 17 Alloys at 1200° and 1350° F. NACA ACR No. 4C22, 1944.\n\n3. Freeman, J. W., Reynolds, E. E., and White, A. E.: High Temperature Alloys Developed for Aircraft Turbosuperchargers and Gas Turbines. Symposium on Materials for Gas Turbines, A.S.T.M., 1946.\n\n4. Freeman, J. W., and Cross, H. C.: A Metallurgical Investigation of a Large Forged Disc of Low-Carbon N-155 Alloy. NACA ARR No. 5K20, 1945.\n\n5. Cross, Howard C., and Freeman, J. W.: A Metallurgical Investigation of Large Forged Discs of Low-Carbon N-155 Alloy. NACA TN No. 1230, 1947.", "timestamp": "2026-07-22T04:58:59.807237+00:00"}
{"citation_id": "19930082498", "source_url": "https://ntrs.nasa.gov/api/citations/19930082498/downloads/19930082498.pdf", "page_number": 21, "total_pages": 49, "image_filename": "19930082498_p21.jpg", "text": "```markdown\nTABLE II.- RESULTS OF MUFFLER INVESTIGATION MADE WITH PROPELLER REMOVED - Continued\n\n| Muffler configuration | Engine speed (rpm) | Fundamental firing frequency, F (cps) | Sound-pressure level (db) | | | | | | | | | | | | | | | Other sounds | | | Back pressure |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | | Over-all | 0.3F | 1.0F | 1.3F | 2.0F | 2.3F | 3.0F | 3.3F | 4.0F | 4.3F | 5.0F | 5.3F | 6.0F | 6.3F | 7.0F | cps | cps | db | |\n| (a) | | | | | | | | | | | | | | | | | | | | | |\n| [Figure: Diagram for row 49] | 2000 | 100.0 | 92.0 | 60 | 92 | 60 | 72 | 72 | 82 | (b) | (b) | (b) | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| | 2790 | 139.5 | 97.0 | 65 | 97 | 75 | 87 | 65 | 81 | 60 | 65 | 70 | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | |\n| [Figure: Diagram for row 50] | 1690 | 82.5 | 79.0 | 45 | 77 | 52 | 73 | 50 | 54 | 52 | 53 | 50 | 50 | 47 | 50 | 51 | 50 | 619 | 50 | 55 | 55 | Medium |\n| | 2000 | 100.0 | 82.5 | 45 | 68 | 50 | 69 | 55 | 58 | 47 | 50 | 47 | 50 | 50 | 50 | 53 | 50 | 750 | 50 | -- | -- | |\n| | 2790 | 139.5 | 87.5 | 53 | 87 | 54 | 83 | 64 | 66 | 57 | 63 | 62 | 68 | 72 | 70 | (b) | (b) | --- | --- | --- | --- | |\n| [Figure: Diagram for row 51] | 1690 | 82.5 | 93.0 | -- | 64 | 87 | 72 | 81 | 57 | 57 | 61 | 65 | 57 | 59 | 52 | 57 | 57 | 59 | 135 | 76 | 370 | 63 | Low |\n| | 2000 | 100.0 | 88.0 | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | |\n| | 2790 | 139.5 | 99.0 | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | |\n| 52 Muffler 51 reversed | 2000 | 100.0 | 87.5 | 46 | 87 | 54 | 74 | 61 | 81 | 64 | 70 | 59 | 54 | 51 | 59 | 61 | 62 | 135 | 68 | 370 | 68 | Low |\n| [Figure: Diagram for row 53] | 2000 | 100.0 | 92.5 | (b) | 92 | (b) | 72 | (b) | 60 | (b) | (b) | (b) | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Low |\n| 54 Muffler 53 reversed | 2000 | 100.0 | 89.5 | (b) | 88 | (b) | 74 | (b) | (b) | (b) | 73 | (b) | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| [Figure: Diagram for row 55] | 1690 | 82.5 | 90.5 | 60 | 88 | 63 | 65 | 62 | 65 | 60 | 65 | 60 | 60 | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| | 2000 | 100.0 | 87.5 | 55 | 87 | 57 | 65 | 59 | 68 | 65 | 63 | 65 | 55 | 50 | 50 | (b) | (b) | 335 | 67 | 435 | 56 | |\n| | 2790 | 139.5 | 88.5 | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | -- | --- | --- | --- | --- | |\n| 56 Muffler 55 reversed | 1690 | 82.5 | 92.0 | 50 | 87 | 55 | 65 | 50 | 63 | 60 | 58 | 60 | (b) | (b) | (b) | (b) | (b) | --- | --- | --- | --- | Medium |\n| | 2000 | 100.0 | 89.0 | 55 | 89 | 60 | 57 | 61 | 78 | 67 | 58 | 59 | 52 | 50 | 50 | (b) | (b) | --- | --- | --- | --- | |\n\n$^a$In the sketches of the configurations, all dimensions are in inches and all cross sections are circular, except where otherwise indicated.\n$^b$The sound-pressure level was below the range of the analyzer.\n\nNACA\nNACA TN No. 1838\n20\n```", "timestamp": "2026-07-22T04:59:01.730198+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 18, "total_pages": 78, "image_filename": "19930082618_p18.jpg", "text": "16\nNACA TN 1945\n\nCONCLUSIONS\n\nFrom investigations of the two-dimensional aerodynamic characteristics of 10 NACA 6-series and 5 NACA 4- and 5-digit-series airfoil sections at seven Reynolds numbers from $9.0 \\times 10^6$ to $0.7 \\times 10^6$, the following conclusions can be drawn:\n\n1. The drag coefficient at the design lift coefficient (designated minimum drag coefficient) of each of the 15 airfoils both in the smooth and rough surface conditions increased as the Reynolds number was lowered from $9.0 \\times 10^6$ to $0.7 \\times 10^6$. The magnitude of this increase became larger for the smooth NACA 6-series sections with increasing airfoil thickness and rearward movement of the position of minimum pressure on the basic thickness form at zero lift. In the rough surface condition and at the lower Reynolds numbers in the smooth surface condition, the reduction in minimum drag to be derived from the use of NACA 6-series as compared with NACA 5-digit-series sections disappeared.\n\n2. Reductions in the Reynolds number generally caused some increase in the extent of the low-drag range for the smooth NACA 6-series airfoils. For all the airfoils, the actual extent of the low-drag range was greater than the theoretical value at the lower Reynolds numbers.\n\n3. Decreasing the Reynolds number from $9.0 \\times 10^6$ to $0.7 \\times 10^6$ caused reductions in the maximum lift of all the airfoils with and without split flaps, in both the smooth and rough surface conditions. The magnitude and character of this reduction, however, varied rather inconsistently with airfoil design and surface condition so that the comparative merits of the group of airfoils changed markedly and in a rather unpredictable manner with Reynolds number and surface conditions.\n\n4. In general, reductions in the Reynolds number appeared to decrease the sharpness of the stall on those NACA 6-series airfoils for which the lift curves are characterized by rather abrupt losses in lift at the stall. The very undesirable stalling characteristics of the NACA 230-series sections were not improved in either the smooth or rough surface condition by reductions in the Reynolds number.\n\n5. Some decrease in the lift-curve slopes of the smooth and rough airfoils accompanied decreases in the Reynolds number. The type and magnitude of the scale effect changed to a small degree, with variations in the airfoil-design parameters considered. In most cases the angle of zero lift seemed to be almost independent of variations in the Reynolds number.", "timestamp": "2026-07-22T04:59:02.653237+00:00"}
{"citation_id": "19930085487", "source_url": "https://ntrs.nasa.gov/api/citations/19930085487/downloads/19930085487.pdf", "page_number": 9, "total_pages": 36, "image_filename": "19930085487_p9.jpg", "text": "NACA RM No. E8J22\n7\n\nTABLE II - NATURAL FREQUENCIES OF VARIOUS\nMODES OF BLADE VIBRATION\n\n| Stage | Blade | Natural frequencies (cps) | | |\n| :--- | :--- | :--- | :--- | :--- |\n| | | **Vibration mode** | | |\n| | | First bending | Second bending | First torsional |\n| 1 | 7 | 394 | 1940 | 2820 |\n| 2 | 7 | 486 | 2520 | 3000 |\n| 3 | 8 | 514 | 2640 | 3240 |\n| 4 | 8 | 626 | 3380 | 3660 |\n| 5 | 8 | 750 | 3780 | 3900 |\n| 6 | 10 | 778 | 3920 | 4340 |\n| 7 | 10 | 900 | 4440 | 4780 |\n| 8 | 10 | 1120 | 5660 | 5180 |\n| 9 | 18 | 1154 | 5420 | 6160 |\n| 10 | 18 | 1220 | 5000 | 6920 |\n\n[Figure: NACA logo]", "timestamp": "2026-07-22T04:59:02.981003+00:00"}
{"citation_id": "19930082614", "source_url": "https://ntrs.nasa.gov/api/citations/19930082614/downloads/19930082614.pdf", "page_number": 19, "total_pages": 36, "image_filename": "19930082614_p19.jpg", "text": "NACA TN 1939\n\nK = 0.0644 × 10⁻³ per foot\n\nV = $\\frac{15,500}{t + 22.2}$\n\nThe relations for speed as a function of time are shown in figure 9 for level flight. Figure 9 indicates that, by reducing the altitude from 25,000 to 10,000 feet, the time required for a given speed reduction is decreased by 40 percent.\n\nConstant Dive Angle\n\nThe variation of speed with time during the first 15 seconds for a constant-angle dive has been calculated for the same assumed coefficients. A dive angle of 60° was assumed.\n\nInitial altitude, 25,000 feet\n\nEstimated average altitude, 20,500 feet\n\nK = 0.0000458 per foot (fig. 1(b))\n\nL = -g sin γ = (32.2) (0.866) = 27.9 feet per second squared\n\n$\\sqrt{L/K} = 780.5$ feet per second\n\nSince $\\sqrt{L/K}$ is greater than V₀, equation (5) is used.\n\nC₂ = $\\frac{1}{2} \\sqrt{K/L} \\log_e \\frac{\\sqrt{L} + V_0 \\sqrt{K}}{\\sqrt{L} - V_0 \\sqrt{K}} = 0.001866$\n\nV = 780.5 tanh [780.5 (0.0000458t + 0.001866)]\n\nInitial altitude, 10,000 feet\n\nAverage altitude, 5,000 feet\n\nK = 0.0000752 per foot\n\n$\\sqrt{L/K} = 609$ feet per second\n\nSince $\\sqrt{L/K}$ is less than V₀, equation (4) is used.\n\nV = 609 coth [609(0.0000752t + 0.002187)]", "timestamp": "2026-07-22T04:59:06.575907+00:00"}
{"citation_id": "19930082485", "source_url": "https://ntrs.nasa.gov/api/citations/19930082485/downloads/19930082485.pdf", "page_number": 37, "total_pages": 62, "image_filename": "19930082485_p37.jpg", "text": "36\nNACA TN No. 1810\n\nWeight-flow parameter, $\\left(\\frac{\\rho_e V_x}{\\rho_t V_{cr}}\\right)_1$\n\nRatio of stagnation-to-ambient-static pressure, $\\frac{P_{t,i}}{P_{s,0}}$\n\n| | |\n| :--- | :--- |\n| O | Cascade tunnel data |\n| $\\Delta$ | Test condition |\n\n[Figure: Graph showing the variation of weight-flow parameter with ratio of stagnation-to-ambient-static pressure. The curve starts near (1.0, 0) and rises, passing through points approximately at (1.1, 0.07), (1.15, 0.13), (1.2, 0.20), (1.3, 0.23), (1.4, 0.245), (1.5, 0.25), (1.6, 0.255), (1.8, 0.26), (2.0, 0.265). Data points are marked with circles (O) and one triangle ($\\Delta$) at approximately (1.6, 0.25).]\n\nFigure 6. - Variation of weight-flow parameter at cascade entrance with ratio of stagnation-to-ambient-static pressure across cascade.\n\n1026", "timestamp": "2026-07-22T04:59:13.968843+00:00"}
{"citation_id": "19930082914", "source_url": "https://ntrs.nasa.gov/api/citations/19930082914/downloads/19930082914.pdf", "page_number": 15, "total_pages": 66, "image_filename": "19930082914_p15.jpg", "text": "14\nNACA TN No. 1857\n\nBecause n is very nearly equal to unity, the equation can be written\n\n$$n - 1 = kp \\tag{1}$$\n\nwhere the constant of proportionality k is the Gladstone-Dale constant.\nThen\n\n$$\\frac{n - 1}{\\rho} = \\frac{n' - 1}{\\rho'}$$\n\nor\n\n$$n' - n = (n - 1)\\left(\\frac{\\rho'}{\\rho} - 1\\right) \\tag{2}$$\n\nNow let one of the two beams pass, for the distance L, through air of a different density $\\rho'$ and index $n'$, and let values of density and index be constant along the length L, but let them be functions of the vertical coordinate y. Then the wavefronts will be distorted, as in figure 9. The amount of distortion, or retardation X at any point is a function of the velocity of light at that point. The time for passage of light of velocity $V'$ through a distance L is $L/V'$. In that same time, light of velocity V will pass through a distance $L + X$. Therefore\n\n$$\\frac{L}{V'} = \\frac{L + X}{V}$$\n\nBut since\n\n$$\\frac{V}{V'} = \\frac{\\lambda}{\\lambda'}$$\n\n$$\\frac{L}{\\lambda'} = \\frac{L + X}{\\lambda}$$", "timestamp": "2026-07-22T04:59:14.767716+00:00"}
{"citation_id": "19930082511", "source_url": "https://ntrs.nasa.gov/api/citations/19930082511/downloads/19930082511.pdf", "page_number": 29, "total_pages": 99, "image_filename": "19930082511_p29.jpg", "text": "NACA TN No. 1826\n27\n\nwhence, by condition (5)\n$$ \\frac{-2y_1}{|z_1|^2} \\frac{A}{a} = -2y_1 C $$\nor\n$$ C = \\frac{A}{a|z_1|^2} \\quad (16) $$\nThe complex equation (15) and the real equation (16) suffice for evaluating the three real constants A, B, and C in equation (14).\n\nTunnel-interference velocity.- The tunnel interference velocity is\n$$ q_1(\\xi, \\xi_1) = i \\left( \\frac{1}{z - z_1} - \\frac{1}{z - \\bar{z}_1} \\right) (A + Bz + Cz^2) \\sqrt{\\frac{1 - z^2}{a^2 - z^2}} + \\frac{i\\Gamma'}{2\\pi(\\xi - \\xi_1)} \\quad (17) $$\nIf the vortex is on the tunnel axis, that is, if $z_1 = iy_1$, equation (15) gives\n$$ \\left. \\begin{aligned} B &= -\\frac{\\Gamma'}{2} \\sqrt{\\frac{a^2 + y_1^2}{1 + y_1^2}} \\\\ C &= \\frac{A}{y_1^2} \\end{aligned} \\right\\} \\quad (18) $$\nComparing this last equation with equation (16), which reduces to the following form for $z_1 = iy_1$\n$$ C = \\frac{A}{ay_1^2} $$\nshows that for this symmetrical case\n$$ A = C = 0 $$", "timestamp": "2026-07-22T04:59:15.341663+00:00"}
{"citation_id": "19930093773", "source_url": "https://ntrs.nasa.gov/api/citations/19930093773/downloads/19930093773.pdf", "page_number": 42, "total_pages": 47, "image_filename": "19930093773_p42.jpg", "text": "NACA RM E9G09\n41\n\n[Figure: Graph showing Engine total-temperature ratio, $T_7/T_1$ vs. Engine total-pressure ratio, $P_7/P_1$]\n\nAltitude (ft)\n○ 5,000\n□ 15,000\n◇ 25,000\n△ 35,000\n▽ 45,000\n◁ 50,000\n\n(a) Flight Mach number, 0.21; altitude, 5000 to 50,000 feet.\nFigure 7. - Variation of engine total-temperature ratio with engine total-pressure ratio.", "timestamp": "2026-07-22T04:59:18.971941+00:00"}
{"citation_id": "19930082617", "source_url": "https://ntrs.nasa.gov/api/citations/19930082617/downloads/19930082617.pdf", "page_number": 20, "total_pages": 58, "image_filename": "19930082617_p20.jpg", "text": "NACA TN 1962\n19\n\nStringers\n$\\circ$ 1 to 9\n$\\times$ 10 to 16\n\nMoment\n(in. - lb)\n1 36.0 X $10^3$\n2 108.0 X $10^3$\n3 180.0 X $10^3$\n4 252.0 X $10^3$\n\n2.57\"\nA\nBand B\nA\nA-A\n\nDistance from horizontal diameter, in.\n10\n9\n8\n7\n6\n5\n4\n3\n2\n1\n0\n1\n2\n3\n4\n5\n6\n7\n8\n9\n10\n\n16 12 8 4 0 -4 -8 -12 -16 -20 X $10^{-4}$\nStrain\n\n1 2 3 4\n\nNACA\n\nFigure 8.- Strain diagram of cylinder 74. Band B.", "timestamp": "2026-07-22T04:59:19.165903+00:00"}
{"citation_id": "19930082450", "source_url": "https://ntrs.nasa.gov/api/citations/19930082450/downloads/19930082450.pdf", "page_number": 26, "total_pages": 37, "image_filename": "19930082450_p26.jpg", "text": "```markdown\nNACA TN No. 1778\n25\n\n<!-- Image (193, 110, 836, 909) -->\n\nFigure 4. -Concluded. $\\frac{t_w}{t_s} = 0.79$.\n```", "timestamp": "2026-07-22T04:59:23.765134+00:00"}
{"citation_id": "19930082476", "source_url": "https://ntrs.nasa.gov/api/citations/19930082476/downloads/19930082476.pdf", "page_number": 33, "total_pages": 41, "image_filename": "19930082476_p33.jpg", "text": "NACA TN No. 1801\n31\n\n$$\n\\frac{I_Z - I_X}{mb^2}\n$$\nRelative mass distribution\nincreased along the fuselage\n\n$$\n\\frac{I_Y - I_Z}{mb^2}\n$$\nRelative mass distribution\nincreased along the wings\n\n[Figure: A chart with a diagonal axis labeled \"$\\frac{I_X - I_Y}{mb^2}$\". The top-left region is labeled \"Loading chiefly along the fuselage\" with values -200 and -100. The bottom-right region is labeled \"Loading chiefly along the wings\" with values 0 and $100 \\times 10^{-4}$. The chart contains a grid and four circled data points labeled 1, 2, 3, and 4. The NACA logo is present in the bottom right corner of the chart area.]\n\nFigure 4.- Mass parameters for loadings tested on the model.", "timestamp": "2026-07-22T04:59:24.051023+00:00"}
{"citation_id": "19930082712", "source_url": "https://ntrs.nasa.gov/api/citations/19930082712/downloads/19930082712.pdf", "page_number": 13, "total_pages": 14, "image_filename": "19930082712_p13.jpg", "text": "```markdown\nNACA TN 1998\n\nR\n$\\Delta$ 6.0 x $10^6$\n$\\nabla$ 9.0\n$\\triangleright$ 11.0\nFlagged symbols denote\nstandard roughness\n\nR\n$\\circ$ 1.8 x $10^6$\n$\\square$ 2.6\n$\\diamond$ 3.0\nFlagged symbols denote\nstandard roughness\n\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\n\nSection drag coefficient, $c_d$\nSection lift coefficient, $c_l$\n\nR\n$\\circ$ 1.8 x $10^6$\n$\\square$ 2.6\n$\\diamond$ 3.0\n$\\triangle$ 6.0\n$\\nabla$ 9.0\n$\\triangleright$ 11.0\n\nR\n$\\circ$ 1.8 x $10^6$\n$\\square$ 2.6\n$\\diamond$ 3.0\n$\\triangle$ 6.0\n$\\nabla$ 9.0\n$\\triangleright$ 11.0\n\na.c. position\nx/c y/c\nSmooth airfoil\n0.284 -0.007\n.275 .021\n.274 .025\n.274 .025\n.274 .024\n.271 .025\n\nRough airfoil\n.282 -.006\n.278 -.007\n.268 .023\n.265 .022\n.254 .024\n.260 .025\n\nMoment coefficient, $c_m$\nSection lift coefficient, $c_l$\n\nMoment coefficient, $c_m$\nSection lift coefficient, $c_l$\n\nNACA\n\n(b) Section pitching-moment characteristics about the aerodynamic center and section\ndrag characteristics.\n\nFigure 1.- Concluded.\n\n11\n\n0012-64\n```", "timestamp": "2026-07-22T04:59:27.490596+00:00"}
{"citation_id": "19930085542", "source_url": "https://ntrs.nasa.gov/api/citations/19930085542/downloads/19930085542.pdf", "page_number": 3, "total_pages": 46, "image_filename": "19930085542_p3.jpg", "text": "1E\nNACA RM No. L8L29\n\nNATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nRESEARCH MEMORANDUM\n\nLOW-SPEED STATIC-STABILITY AND ROLLING CHARACTERISTICS OF\nLOW-ASPECT-RATIO WINGS OF TRIANGULAR AND\nMODIFIED TRIANGULAR PLAN FORMS\n\nBy Byron M. Jaquet and Jack D. Brewer\n\nSUMMARY\n\nA low-speed investigation was made in the Langley stability tunnel\nto determine experimentally the effects of changes in profile and aspect\nratio on the low-speed static-stability and rolling characteristics of\ntriangular wings. The investigation was extended to determine the effects\nof adding fins to the upper surface and of cutting portions from the tips\nof a triangular wing to form low-aspect-ratio tapered wings.\n\nIn general, profile had little effect on the static-stability and\nrolling characteristics of triangular wings at low lift coefficients.\nThe greatest effect of profile was in the high-lift-coefficient range.\nThe linear range of the static-stability parameters was decreased (in\nmuch the same manner as for untapered swept wings) as the leading-edge\nsharpness was increased.\n\nSeveral of the characteristics of triangular wings may be estimated\nwith fair accuracy by available swept-wing theory. Available low-aspect-\nratio triangular-wing theory was found to be reliable for certain charac-\nteristics, but for others, particularly the lift-curve slope, aerodynamic\ncenter position, and the damping in roll, the agreement was poor except at\nthe lowest test aspect ratio (A = 1.07).\n\nThe vertical fins tested provided good directional stability through-\nout the entire lift-coefficient range. The fins increased the damping\nin roll and decreased the variation of the effective dihedral parameter\nwith lift coefficient.\n\nThe series of modified triangular wings obtained by cutting various\nportions from the tips of a basic triangular wing generally had good\nlongitudinal and directional stability but very high effective dihedral.\nMost of the characteristics of these wings, at low lift coefficients,\ncould be predicted with fair accuracy by available swept-wing theory.", "timestamp": "2026-07-22T04:59:29.226481+00:00"}
{"citation_id": "19930085519", "source_url": "https://ntrs.nasa.gov/api/citations/19930085519/downloads/19930085519.pdf", "page_number": 9, "total_pages": 46, "image_filename": "19930085519_p9.jpg", "text": "```markdown\n8\nNACA RM No. L8K19\n\nconditions of the wing model and not by the change in Reynolds numbers\nsince the Zap flap test and the accompanying plain-wing test were per-\nformed before the installation of the slotted flap, the plug ailerons,\nand the various plain-spoiler configurations reported in reference 1. It\nis believed that the maximum lift coefficients for the full-span and\npartial-span slotted flaps would have been somewhat higher had the slotted-\nflap tests been performed with the wing in a condition comparable to that\nfor the Zap flap tests.\n\nFigure 14 presents the trimmed-gliding characteristics of the plain\n42° sweptback wing and the wing equipped with various flap configurations.\nThe gliding characteristics were calculated for an airplane having an\nassumed wing loading of 40 pounds per square foot and a tail length of\n3.05. (The gliding characteristics of the wing with the half-span split\nflap, presented in fig. 14, were calculated from unpublished data\nobtained in the Langley 19-foot pressure tunnel.) At a sinking speed\n$V_S$ of 30 feet per second (assumed to be the maximum permissible) the\nplain wing had the highest gliding speed, about 135 miles per hour.\n\nAt a sinking speed of 30 feet per second the gliding speed decreased\nto about 132 miles per hour with either the half-span Zap flap or the\nhalf-span split flap deflected and decreased to about 118 miles per hour\nwith the half-span slotted flap deflected 50°. The slowest landing speeds\n(about 110 mph) were obtained with the full-span slotted flap deflected\neither 30° or 50°. On the basis of these results the full-span slotted\nflap at 30° deflection is considered to be the most satisfactory flap con-\nfiguration for this particular wing. The data presented in figure 14 are\nfor relatively low Reynolds numbers. The unpublished data from the\nLangley 19-foot pressure tunnel indicate that increasing Reynolds numbers\nresult in an increase in the value of maximum lift coefficient which\nwould, of course, result in a somewhat lower landing speed for each of the\nflap configurations.\n\nTuft studies of the flow along the wing lower surface in the vicinity\nof the flap slot with the full-span slotted flap deflected showed a large\namount of spanwise air flow toward the wing tip in lines approximately\nparallel to the flap leading edge. The flap-slot flow-control vanes A\nand B shown in figures 5 and 6, respectively, were therefore installed\non the wing lower surface in the flap slot to interrupt this spanwise flow\nand to direct it in lines perpendicular to the flap leading edge, thus\nincreasing the dynamic pressure and, consequently, the lift over the flap.\nThe results presented in figure 15 for the wing with tufts and the full-\nspan slotted flap deflected 50° show that the wing lift was decreased and\nthe wing drag increased by the installation of vanes A. The installation\nof flow-control vanes B on the wing with the full-span slotted flap\ndeflected 50° resulted in an increase in both the wing lift and drag as\nshown in figure 16. Comparison of the calculated gliding characteristics\n```", "timestamp": "2026-07-22T04:59:29.913620+00:00"}
{"citation_id": "19930082585", "source_url": "https://ntrs.nasa.gov/api/citations/19930082585/downloads/19930082585.pdf", "page_number": 21, "total_pages": 30, "image_filename": "19930082585_p21.jpg", "text": "```markdown\n20\nNACA TN 1907\n\n<!-- Image (103, 101, 817, 764) -->\n\nFigure 1.- Effect of rate of pitch reduction on the variations of descending velocity and rotor angular velocity with time after power failure. Flapping neglected except as noted. $I_1 = 200$ slug-feet$^2$.\n```", "timestamp": "2026-07-22T04:59:32.429071+00:00"}
{"citation_id": "19930085471", "source_url": "https://ntrs.nasa.gov/api/citations/19930085471/downloads/19930085471.pdf", "page_number": 11, "total_pages": 28, "image_filename": "19930085471_p11.jpg", "text": "```markdown\nTABLE I\n\nEXPERIMENTAL AND THEORETICAL RESULTS OF FLUTTER INVESTIGATIONS\n\n| Parameters | Model designation | A-1 | B-1 | B-2 | B-3 | B-4 | B-5 | C-1 | C-2 | D-1 | E-1 | F-1 | G-1 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Description of model section | | NACA 65-007 | NACA 16-010 | NACA 16-010 | NACA 16-010 | NACA 16-010 | NACA 16-010 | 8 percent circular arc | 8 percent circular arc | Modified circular arc 4.74 percent thick | 3 percent circular arc | 3 percent double wedge | Modified circular arc 5 percent thick |\n| Length, in. | | 6.06 | 9.125 | 9.125 | 9.125 | 9.125 | 7.50 | 6.00 | 6.00 | 7.125 | 9.125 | 9.125 | 6.00 |\n| Chord, in. | | 3.03 | 4.03 | 4.03 | 4.03 | 4.03 | 4.03 | 3.03 | 3.03 | 4.22 | 4.04 | 4.01 | 4.00 |\n| Center of gravity, percent chord | | 49.1 | 51.6 | 54.6 | 56.0 | 56.7 | 57.0 | 53.0 | 55.4 | 46.0 | 62.6 | 56.5 | 53.5 |\n| Elastic axis, percent chord | | 41.3 | 34.1 | 39.6 | 39.6 | 44.2 | 39.5 | 48.0 | 51.55 | 37.0 | 38.7 | 45.2 | 47.5 |\n| $r_\\alpha^2$ | | 0.28 | 0.39 | 0.38 | 0.40 | 0.37 | 0.37 | 0.230 | 0.233 | 0.275 | 0.510 | 0.29 | 0.27 |\n| $1/\\mu$ | | 64.9 | 95.3 | 108.1 | 113.1 | 113.3 | 130 | 67.1 | 74.1 | 53.5 | 267.5 | 150.8 | 51.7 |\n| $\\Delta J$ | | 578 | 3060 | 2945 | 3150 | 4105 | 2500 | 839 | 909.5 | 1025 | 2710 | 2220 | 1066 |\n| $f_h$, cps | | 133.4 | 95.0 | 100.9 | 98.0 | 104.2 | 104.0 | 157 | 157 | 110 | 25.2 | 28.5 | 148.5 |\n| $f_\\alpha$, cps | | 278 | 163 | 156.3 | 154.8 | 183 | 162 | 363 | 361 | 178 | 81.7 | 132.8 | 245 |\n| $f_f$, cps | | 180 | 134 | 132.5 | 134.6 | 131.5 | 136 | 188 | 184 | 142 | 71 | 70.5 | 176 |\n| $\\rho \\times 10^3$ (test section) | | 0.876 | 0.945 | 0.890 | 0.897 | 0.908 | 0.886 | 0.887 | 0.884 | 0.918 | 0.888 | 0.888 | 0.910 |\n| $\\delta_h$ | | 0.04 | 0.03 | 0.025 | 0.03 | 0.025 | 0.035 | 0.05 | 0.05 | 0.03 | 0.004 | 0.01 | 0.04 |\n| $\\delta_\\alpha$ | | 0.04 | 0.04 | 0.04 | 0.035 | 0.03 | 0.035 | 0.04 | 0.04 | 0.03 | 0.005 | 0.005 | 0.04 |\n| $\\omega_h/\\omega_\\alpha$ | | 0.48 | 0.583 | 0.645 | 0.633 | 0.57 | 0.64 | 0.432 | 0.435 | 0.614 | 0.308 | 0.215 | 0.606 |\n| $V/b\\omega_\\alpha$ (experimental) | | 10.15 | 9.98 | 10.31 | 10.20 | 10.40 | 10.03 | 9.725 | 9.92 | 9.04 | 19.13 | 19.61 | 7.71 |\n| $V/b\\omega_\\alpha$ (theoretical) | | 6.946 | 7.770 | 7.109 | 7.546 | 7.938 | 7.683 | 8.53 | 7.18 | 5.28 | 14.85 | 9.38 | 4.92 |\n| $V/b\\omega_\\alpha$ (experimental) | | 6.59 | 8.21 | 8.74 | 8.91 | 7.44 | 8.42 | 5.02 | 5.055 | 7.25 | 16.7 | 10.35 | 5.54 |\n| $V/b\\omega_\\alpha$ (theoretical) | | 6.554 | 7.158 | 5.664 | 6.029 | 6.234 | 6.041 | 4.513 | 5.503 | 5.069 | 13.08 | 8.23 | 3.905 |\n| $\\omega_f/\\omega_\\alpha$ (experimental) | | 0.648 | 0.822 | 0.847 | 0.870 | 0.719 | 0.840 | 0.518 | 0.51 | 0.798 | 0.868 | 0.531 | 0.718 |\n| $\\omega_f/\\omega_\\alpha$ (theoretical) | | 0.944 | 0.921 | 0.797 | 0.800 | 0.785 | 0.786 | 0.577 | 0.766 | 0.96 | 0.88 | 0.868 | 0.795 |\n| $(V/b\\omega_\\alpha)_{M=0}$ (theoretical) | | 3.427 | 4.302 | 4.245 | 4.358 | 4.121 | 4.612 | 3.184 | 3.374 | 2.811 | 8.263 | 5.491 | 2.585 |\n\nNACA\n\nNACA RM No. L8J11\n\n[Stamp: UNCLASSIFIED]\n[Stamp: CONFIDENTIAL]\n[Stamp: UNCLASSIFIED]\n[Stamp: CONFIDENTIAL]\n\n9\n```", "timestamp": "2026-07-22T04:59:37.772484+00:00"}
{"citation_id": "19930082618", "source_url": "https://ntrs.nasa.gov/api/citations/19930082618/downloads/19930082618.pdf", "page_number": 19, "total_pages": 78, "image_filename": "19930082618_p19.jpg", "text": "NACA TN 1945\n17\n\n6. The value of the quarter-chord pitching moment at the design\nangle of attack did not vary with Reynolds number for the plain airfoils.\nThe chordwise position of the aerodynamic center varied somewhat with\nReynolds number, but these variations were, in most cases, relatively\nsmall. With 0.20-chord simulated split flaps deflected 60°, the value\nof the quarter-chord pitching moment at zero angle of attack in many\ncases varied somewhat in magnitude with Reynolds number.\n\nLangley Aeronautical Laboratory\nNational Advisory Committee for Aeronautics\nLangley Air Force Base, Va., July 6, 1949\n\nREFERENCES\n\n1. Abbott, Ira H., Von Doenhoff, Albert E., and Stivers, Louis S., Jr.:\nSummary of Airfoil Data. NACA Rep. 824, 1945.\n\n2. Loftin, Laurence K., Jr., and Cohen, Kenneth S.: Aerodynamic\nCharacteristics of a Number of Modified NACA Four-Digit-Series\nAirfoil Sections. NACA TN 1591, 1948.\n\n3. Loftin, Laurence K., Jr.: Theoretical and Experimental Data for a\nNumber of NACA 6A-Series Airfoil Sections. NACA TN 1368, 1947.\n\n4. Von Doenhoff, Albert E., and Abbott, Frank T., Jr.: The Langley Two-\nDimensional Low-Turbulence Pressure Tunnel. NACA TN 1283, 1947.\n\n5. Loftin, Laurence K., Jr., and Bursnall, William J.: The Effects of\nVariations in Reynolds Number between $3.0 \\times 10^6$ and $25.0 \\times 10^6$\nupon the Aerodynamic Characteristics of a Number of NACA 6-Series\nAirfoil Sections. NACA TN 1773, 1948.", "timestamp": "2026-07-22T04:59:41.427617+00:00"}