PDF/ocr_dataset_102.jsonl

{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 15, "total_pages": 24, "image_filename": "19930091716_p15.jpg", "text": "# NEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS\n\n## CONSIDERATIONS IN APPLYING NEGATIVE THRUST AND TORQUE DATA TO THE SOLUTION OF PROBLEMS\n\n**Coefficients.**—The thrust coefficient $T_c$ is especially suited to a negative-thrust analysis because it does not involve the engine speed and because of its similarity to the usual drag coefficient. It is not convenient to use the normal propeller thrust coefficient $C_T=T_c/\\rho n^2 D^4$ because $C_T$ approaches infinity as $nD/V$ approaches zero, and difficulty in plotting arises. With the diameter and velocity known, the thrust may be easily calculated for any value of $T_c$.\n\n**Effect of engine.**—In most problems involving negative thrust, the propeller is mounted on an airplane engine, which may be “dead” (switch off and being turned over against its own friction), partly throttled, or operating at full throttle. The main difficulty in calculating the negative thrust of a propeller operating under any given condition, especially the one where it is turning a dead engine, is found in determining the engine speed. In the special case of the “freewheeling” propeller, the thrust coefficient is easily found, it being the value corresponding to the $nD/V$ where the torque coefficient is zero. When the propeller is turning a dead engine, however, the revolution speed depends upon the friction torque of the engine, which is itself an extremely variable quantity.\n\n**The coefficient $Q_e$.**—In reference 5 it was pointed out that flight tests indicated the rotational speeds of propellers turning dead engines on multiengine airplanes to be from 35 to 50 percent of rated engine speed. It was also pointed out that, through this range of engine speeds, the friction torque of the average airplane engine might be represented by an equation of the following type:\n\n$$Q_f=k_1 N_e \\Delta(1+k_2 h)/k_3 G. R.$$\n\nwhere $\\Delta$ is the engine displacement, cu. in.  \n$N_e$, crankshaft revolution speed, r. p. m.  \n$G. R.$, the ratio of propeller speed to crankshaft speed.  \n$k_1$, $k_2$, and $k_3$, appropriate constants.\n\nFor a particular engine and at a given altitude $h$, the equation is simplified to:\n\n$$\\frac{Q_f}{N_e}=\\text{constant } (K)$$\n\nThis approximation was used in reference 5 to develop the following form of coefficient:\n\n$$Q_e=\\frac{Q}{\\rho V^2 D^3} \\times \\frac{V}{nD} = \\frac{Q/n}{\\rho V D^3}$$\n\nThis relation may be put in a more useful form:\n\n$$Q_e=17,200 \\times \\frac{Q_f/N}{\\sigma V_1 D^3}$$\n\nwhere $Q_f$ is engine friction torque (considered negative), ft.-lb.  \n$N$, propeller revolution speed, r. p. m.  \n$V_1$, air speed, m. p. h.  \n$\\sigma$, relative density, $\\rho/\\rho_0$.  \n$\\rho_0$, mass density of the air at sea level, slugs per cu. ft.\n\nIn references 3 and 5, charts are presented having thrust coefficient $T_c$ plotted against $Q_e$. Thus, when the value of $Q_f/n$ is known for the engine in question, the value of $Q_e$, at any given altitude and for any given propeller diameter, will depend only on the velocity. The plot of $T_c$ against $Q_e$ then becomes, for any particular case, a plot of the thrust against the inverse of the velocity. It appears likely, from the friction data shown in references 3 and 5, that the usefulness of the relation $Q_f/N_e=K$ will extend over a greater range of engine speeds than previously indicated.\n\n**The $Q_e$ modification to negative-thrust charts.**—In the present report it is shown that a slight modification of the usual plots of $T_c$ and $Q_e$ against $nD/V$ will provide the equivalent of a plot of $T_c$ against $Q_e$, such as given in references 3 and 5.\n\nThe necessary modification to the usual charts is explained as follows: Since $Q_e=\\frac{Q_e}{nD/V^2}$ it is clear that the locus of all points having a single value of $Q_e$ may be represented on plots of $Q_e$ against $nD/V$ (figs. 3 to 9) by a straight line passing through the origin. The position of this line for any particular value of $Q_e$ is easily determined on the chart from the fact that $Q_e=Q_e$ where $\\frac{nD}{V}=1$. Although the position of this line is easily determined, a scale, or rather a director line, has been placed on each chart showing the intersections of the $Q_e$ lines for various values of $Q_e$.\n\nIn actual use a straightedge placed from the origin to the desired value of $Q_e$ on the scale will permit values of the coefficients to be read without drawing the line.\n\n**Examples of use of $Q_e$.**—As an example, suppose the friction torque of the engine and the velocity are such as to make the value of $Q_e=-0.0019$ and that it is desired to find the value of $T_c$ at a blade angle of $20^\\circ$. The solution of this problem is indicated in figure 7. The broken line O–C represents a constant value of $Q_e$. Where this line intersects the $20^\\circ$ blade-angle curve at point D, project up along the line D–D′′ to the point D′′. The $T_c$ coordinate of the point D′′ is the desired value.\n\nThe propeller speed may be calculated from the value of $nD/V$ at point D′ as follows:\n\n$$N=\\frac{nD}{V} \\frac{V_1 \\times 88}{D}$$", "timestamp": "2026-07-19T18:19:14.667676+00:00"}
{"citation_id": "19930091715", "source_url": "https://ntrs.nasa.gov/api/citations/19930091715/downloads/19930091715.pdf", "page_number": 27, "total_pages": 30, "image_filename": "19930091715_p27.jpg", "text": "CHARACTERISTICS OF FULL-SCALE PROPELLERS HAVING 2, 3, AND 4 BLADES 23\n\n5. Compute $T_0 = [\\eta_0 \\times (\\text{b. hp.})_0 \\times 375] / V_0$\n\n6. Compute $V = V_0 \\times \\frac{J}{J_0} \\times \\frac{N}{N_0}$\n\n7. Compute thrust, $T = T_0 \\times \\frac{C_{T_0}}{C_{T_0}} \\times \\frac{C_T}{C_P} \\times K \\times \\frac{C_T}{C_P}$\n\nwhere $K = \\frac{T_0 C_{T_0}}{C_{T_0}}$\n\nThis method assumes that the full-throttle engine torque is constant.\n\nAs an example, assume that it is desired to obtain the propeller thrust through the take-off and climbing ranges for an airplane having the following characteristics:\n\n$V_0 = 190$; $N_0 = 1,500$; $(\\text{b.hp.})_0 = 600$; $J_0 = 1.00$; $D_0$(2-blade) = 11 feet 1½ inches; $\\eta_0 = 0.862$; $\\beta_0 = 25^\\circ$.\n\nBlade section = Clark Y.\n\nThe computed data may be conveniently tabulated as follows:\n\n$C_{T_0} = 0.0448$; $C_{P_0} = 0.0520$; $T_0 = 1,020$ lb.; $J_0 = 1.00$.\n\n| J | J/J₀ | Cₜ | Cₚ | Cₜ/Cₚ | Cₚ₀/Cₚ | N/N₀ | V (m. p. h.) | T (lb.) |\n|---|------|-----|-----|--------|---------|-------|--------------|---------|\n| 0.1 | 0.1 | 0.110 | 0.1056 | 1.042 | 0.493 | 0.792 | 13.3 | 1.232 |\n| .2 | .2 | .1405 | .1017 | 1.059 | .512 | .720 | 27.4 | 1.252 |\n| .3 | .3 | .1058 | .0972 | 1.088 | .535 | .721 | 41.7 | 1.287 |\n| .4 | .4 | .1035 | .0911 | 1.158 | .571 | .755 | 57.4 | 1.370 |\n| .5 | .5 | .1037 | .0838 | 1.237 | .621 | .778 | 73.0 | 1.457 |\n| .6 | .6 | .0970 | .0823 | 1.178 | .632 | .795 | 90.6 | 1.392 |\n| .7 | .7 | .0870 | .0740 | 1.100 | .658 | .811 | 106.0 | 1.300 |\n| .8 | .8 | .0748 | .0732 | 1.022 | .710 | .842 | 128.0 | 1.210 |\n\n$K = \\frac{0.0520}{0.0448} \\times 1,020 = 1,182$\n\nEffect of blade width and body.—The two methods given of calculating thrust, and also the method of selecting propellers, assumed that the propellers under consideration had the same blade width as the ones for which the data are given in this report. Frequently it may be required to find the diameter, the design blade-angle setting, and the thrust of a propeller having a blade width slightly different from those tested. As was mentioned earlier, it may be assumed that the power and the thrust vary directly with the blade areas (or blade widths) for propellers with similar shape characteristics where the differences in areas are small.\n\nIn the calculation of $C_T$, the power should therefore be multiplied by the ratio of the blade widths $b_1/b_2$, where $b_1$ is the blade width at three-quarters radius of the propeller for which the design charts were made and $b_2$ is the blade width at the same radius for the propeller under consideration.\n\nThe same ratio should be used in calculating the value of $C_P$ to be used in obtaining the take-off thrust, and the take-off thrust obtained from the charts should be divided by this ratio to obtain the actual thrust of the propeller.\n\nSimilarly, corrections are necessary in the case where the body under consideration is greatly different from the liquid-cooled engine nacelle with which the present tests were made. Some information with regard to the effect of the body on the propulsive efficiency may be obtained from reference 2. The added drag of those parts of the airplane in the slipstream (other than the body itself) should also be considered. Parts of the wing, the tail surfaces, and the landing gear are often in the propeller slipstream and their added drag due to the slipstream may be approximated from the following relation:\n\n$$\n\\frac{\\Delta D}{D} = 2.5 C_p / J^2\n$$\n\nwhere $\\Delta D$ is the added drag and $D$ is the drag without slipstream.\n\nThe test data for a propeller having one airfoil section should not be used to calculate the performance of a propeller having another airfoil section.\n\nIt is shown in reference 3 that compressibility often has a marked effect on the performance of a propeller in the take-off range. The necessary corrections for compressibility are not easily applied but methods of making such corrections are explained in reference 3.\n\nREFERENCES\n\n1. Weick, Fred E., and Wood, Donald H.: The Twenty-Foot Propeller Research Tunnel of the National Advisory Committee for Aeronautics. T. R. No. 300, N. A. C. A., 1928.\n2. Biermann, David, and Hartman, Edwin P.: Tests of Five Full-Scale Propellers in the Presence of a Radial and a Liquid-Cooled Engine Nacelle, Including Tests of Two Spinners. T. R. No. 642. N. A. C. A., 1938.\n3. Biermann, David, and Hartman, Edwin P.: The Effect of Compressibility of Eight Full-Scale Propellers Operating in the Take-Off and Climbing Range. T. R. No. 639. N. A. C. A., 1938.\n4. Glauert, H.: Airplane Propellers. Vol. IV, div. L of Aerodynamic Theory, W. F. Durand, ed., Julius Springer (Berlin), 1935, p. 201.\n5. Weick, Fred E.: Working Charts for the Selection of Aluminum Alloy Propellers of a Standard Form to Operate with Various Aircraft Engines and Bodies. T. R. No. 350, N. A. C. A., 1930.\n\nU. S. GOVERNMENT PRINTING OFFICE: 1939", "timestamp": "2026-07-19T18:19:14.911035+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 5, "total_pages": 18, "image_filename": "19930094551_p5.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 865\n\nThe pressure head was suspended, as already stated, from an all-metal tubing which at the same time served as supporting and pressure-tapping element. The point of suspension was provided directly above the center of gravity of the pressure head. A link connection between pressure head and tubing kept the lever arm and therewith the turning moment about the transverse axis of the pressure head to a minimum.\n\nIn the final version (fig. 2) the link consists of a sphere carried in the head and terminating at the top in a neck H 4 mm thick and 100 mm long. The upper end of H is hollow with a nipple M at the side. The neck is soldered to the metal tubing. A short piece of rubber hose serving as pressure tap, connects nipple N with the pressure-head nipple N₂, which in turn leads within the pressure head to the annular slot. A safety against turning of the spherical pivot about the normal axis of the pressure head prevents the rubber hose from becoming detached. Both the head and the stabilizing ring are of nickel-plated steel. The total length is 625 mm, for a diameter of D = 45 mm. The annular slot for tapping the static pressure lies 3 D aft of the foremost point of the pressure head. Wind-tunnel tests disclosed a 2.4 percent static-pressure difference as compared with the static reading of a calibrated Pitot tube of normal size.\n\nIII. RESULTS OF TESTS\n\nQualitatively, the flight tests revealed that both the pressure head and tubing remain so much more steady as the extensibility is less and the stability of the tubing greater, that is, the greater the restoring forces are which occur as soon as the tubing curve is changed by some outside influence, such as gusts, for instance. The stability of the tubing (apart from the cable stiffness which, however, is not attempted in the interest of windability), is chiefly influenced by the weight of the instrument. It was noticed that all oscillatory motions set up during the measurements originated in the upper part of the tubing exposed to propeller slipstream and wing wake, while the head itself was only indirectly set in vibration. These vibrations are at right angles to the direction of flight. By increasing the weight from 2.4 kg (weight of the old pressure head) to 5.4 kg (weight of experimental model)", "timestamp": "2026-07-19T18:19:25.375289+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 18, "total_pages": 43, "image_filename": "19930094544_p18.jpg", "text": "16 N.A.C.A. Technical Memorandum No. 872\n\nthe \"Akron,\" In contrast to the \"Akron,\" however, lipped tubes are used in the corners, into which the wall plates grip. In the three boom girders of the spacious rings these tubes are made of high-strength steel, while the wall plates are of duralumin.\n\nThrough the so-called efficiency factor one has a comparison of the values of the girders developed. By this is meant the relationship of the buckling load attained, in tons, to the running girder weight in kg/m. This has the dimension km. In figure 38 the efficiency factors of the triangular girders for LZ 127 and LZ 129 are plotted on the girder cross-sections. It is seen that the efficiency factors of the new girders, in comparison with the earlier ones, have increased significantly. It is especially significant in connection with the girders used, that the efficiency factors increase with increasing cross-section. From this it follows, that the structural improvement of lighter girders is particularly difficult. In figure 39 the efficiency factors for the girders developed by the Goodyear-Zeppelin Corporation are shown. In the case of the girders used in the \"Akron,\" made of the American aluminum alloy 17SRT, they lie between 5 and 8. Moreover, they may be brought higher with the use of the high strength alloy 24SRT and with improved forming.\n\nThose developed for the framing of the R 100 are triangular girders, the tubular booms of which show an especially noteworthy development. Figure 40 shows such a tube in formation. The tubes are rolled in spiral form from strips of plate and riveted along the contacting edges. As is evident from figure 41, the boom tubes are joined by means of box-type struts, which are arranged opposed to one another in a manner similar to that used in the previously described development of the girders of the LZ 129, and which have been provided with lightening holes.\n\nThe longitudinal girders in the R 101 are constructed in yet another manner (fig. 42). These longitudinal girders, which likewise are triangular girders, have booms of steel tubing and struts of duralumin tubing. The rectangular panels are cross-braced by means of wire diagonals. The girders have a considerable depth (up to 70 cm). The steel tubes of the booms are not drawn, but are of sheeting bent together.\n\nIn joint design one can distinguish fundamentally two", "timestamp": "2026-07-19T18:19:40.663508+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 7, "total_pages": 17, "image_filename": "19930094535_p7.jpg", "text": "6 N.A.C.A. Technical Memorandum No. 881\n\n| Solution of | | Behavior in liquid | Change in weight after six months |\n| :--- | :--- | :--- | :--- |\n| soda, | 10 percent | Stable | +0.5 percent |\n| \" | 30 \" | \" | +0.5 \" |\n| Acetone, | | Disintegrates | -- |\n| Ether, | | Unstable | +16.5 percent |\n| Benzol, | | \" | +51. \" |\n| Full mixture. | | \" | +12.5 \" |\n| Chlorated hydrocarbons, | | Disintegrates or dissolves | -- |\n| Cyclohexane, | | Dissolves | -- |\n| Ethyl acetate, | | Disintegrates | -- |\n| Nitrous acid, | | Unstable | +8.3 percent (after 4 months) |\n\nPrincipal Properties of the Polymer Mixtures\n\nSpecific weight . . . . . . . . . 1.34\nTensile strength, about . . . . . 600 kg/cm²\nBending strength, about . . . . 1,000 kg/cm²\nCompressive strength, about . . . 800 kg/cm²\nElasticity modulus . . . . . . 32,000 kg/cm²\nImpact bending strength . . up to 450 cm kg/cm²\nHeat resistance according to Martens, about . . . . . . . 60° C.\nExpansion coefficient . . . 78 x 10⁻⁶", "timestamp": "2026-07-19T18:19:51.008869+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 17, "total_pages": 20, "image_filename": "19930094557_p17.jpg", "text": "N.A.C.A. Technical Memorandum No.859\nFigs.5,6\n\nM = 0; $\\alpha = 20^\\circ$\n\n[Figure: Graph showing Rotation of airfoil vs Jet rotation. Two lines: Theoretical (dashed) and Experimental (solid with circles). Axes range from 0 to 3.]\n\nFigure 5.\n\nM = 0; $\\alpha = 25^\\circ$\n\n[Figure: Graph showing Rotation of airfoil vs Jet rotation. Two lines: Theoretical (dashed) and Experimental (solid with circles). Axes range from 0 to 3.]\n\nFigure 6.\n\nFigures 5,6.- Jet rotation substituted for airfoil rotation.", "timestamp": "2026-07-19T18:19:56.141810+00:00"}
{"citation_id": "19930094542", "source_url": "https://ntrs.nasa.gov/api/citations/19930094542/downloads/19930094542.pdf", "page_number": 67, "total_pages": 102, "image_filename": "19930094542_p67.jpg", "text": "N.A.C.A. Technical Memorandum No. 874\nFigs. 43, 44, 47, 77\n\n[Figure: Wing with motor not enclosed in fairing.]\nFigure 43.- Wing with motor not enclosed in fairing.\n\n[Figure: Wing with motor enclosed in fairing.]\nFigure 44.- Wing with motor enclosed in fairing.\n\n[Figure: Bearing pieces for propeller shaft.]\nFigure 47.- Bearing pieces for propeller shaft.\n\n[Figure: Test set-up for the down-wash measurements with dynamic pressure sphere.]\nFigure 77.- Test set-up for the down-wash measurements with dynamic pressure sphere.", "timestamp": "2026-07-19T18:19:58.273281+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 2, "total_pages": 22, "image_filename": "19930091655_p2.jpg", "text": "# AERONAUTIC SYMBOLS\n\n## 1. FUNDAMENTAL AND DERIVED UNITS\n\n| | Symbol | Metric | | English | |\n| :--- | :---: | :--- | :--- | :--- | :--- |\n| | | Unit | Abbreviation | Unit | Abbreviation |\n| Length<br>Time<br>Force | $l$<br>$t$<br>$F$ | meter<br>second<br>weight of 1 kilogram | m<br>s<br>kg | foot (or mile)<br>second (or hour)<br>weight of 1 pound | ft. (or mi.)<br>sec. (or hr.)<br>lb. |\n| Power<br>Speed | $P$<br>$V$ | horsepower (metric)<br>kilometers per hour<br>meters per second | k.p.h.<br>m.p.s. | horsepower<br>miles per hour<br>feet per second | hp.<br>m.p.h.<br>f.p.s. |\n\n## 2. GENERAL SYMBOLS\n\n$W$, Weight $= mg$\n$g$, Standard acceleration of gravity $= 9.80665$ m/s$^2$ or $32.1740$ ft./sec.$^2$\n$m$, Mass $= \\frac{W}{g}$\n$I$, Moment of inertia $= mk^2$. (Indicate axis of radius of gyration $k$ by proper subscript.)\n$\\mu$, Coefficient of viscosity\n$\\nu$, Kinematic viscosity\n$\\rho$, Density (mass per unit volume)\nStandard density of dry air, $0.12497$ kg$\\cdot$m$^{-4}\\cdot$s$^2$ at $15^\\circ$ C. and $760$ mm; or $0.002378$ lb.$\\cdot$ft.$^{-4}\\cdot$sec.$^2$\nSpecific weight of \"standard\" air, $1.2255$ kg/m$^3$ or $0.07651$ lb./cu.ft.\n\n## 3. AERODYNAMIC SYMBOLS\n\n$S$, Area\n$S_e$, Area of wing\n$G$, Gap\n$b$, Span\n$c$, Chord\n$b^2$\n$S$, Aspect ratio\n$V$, True air speed\n$q$, Dynamic pressure $= \\frac{1}{2}\\rho V^2$\n$L$, Lift, absolute coefficient $C_L = \\frac{L}{qS}$\n$D$, Drag, absolute coefficient $C_D = \\frac{D}{qS}$\n$D_p$, Profile drag, absolute coefficient $C_{D_p} = \\frac{D_p}{qS}$\n$D_i$, Induced drag, absolute coefficient $C_{D_i} = \\frac{D_i}{qS}$\n$D_{p_i}$, Parasite drag, absolute coefficient $C_{D_{p_i}} = \\frac{D_{p_i}}{qS}$\n$C$, Cross-wind force, absolute coefficient $C_C = \\frac{C}{qS}$\n$R$, Resultant force\n$i_w$, Angle of setting of wings (relative to thrust line)\n$i_t$, Angle of stabilizer setting (relative to thrust line)\n$Q$, Resultant moment\n$\\Omega$, Resultant angular velocity\n$\\frac{Vl}{\\mu}$, Reynolds Number, where $l$ is a linear dimension (e.g., for a model airfoil 3 in. chord, 100 m.p.h. normal pressure at $15^\\circ$ C., the corresponding number is 234,000; or for a model of 10 cm chord, 40 m.p.s. the corresponding number is 274,000)\n$C_{p}$, Center-of-pressure coefficient (ratio of distance of c.p. from leading edge to chord length)\n$\\alpha$, Angle of attack\n$\\epsilon$, Angle of downwash\n$\\alpha_0$, Angle of attack, infinite aspect ratio\n$\\alpha_i$, Angle of attack, induced\n$\\alpha_a$, Angle of attack, absolute (measured from zero-lift position)\n$\\gamma$, Flight-path angle", "timestamp": "2026-07-19T18:20:07.479227+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 11, "total_pages": 47, "image_filename": "19930093641_p11.jpg", "text": "9\n\nthe stall begins almost simultaneously at the tips and behind the nacelles. For the wing-nacelle arrangement, the two stalled regions unite at an angle of attack of about $12^\\circ$, after which the lift curve (fig. 10) indicates a general stall for the wing. The flat top of the lift curve is generally characteristic of cases in which nacelle interference exists. Tuft observations were not obtained for the extension-shaft arrangements, but it might be expected that the results would be similar to those for the wing alone.\n\nMaximum L/D ratio.- The maximum L/D ratios in table I show the same general trends indicated by the high-speed drag coefficients and clearly demonstrate that the extension shafts only slightly affect the aerodynamic characteristics of the bare-wing model. The maximum L/D value for the bare-wing model is 19.8, compared with 19.6 for the pusher and 19.0 for the tractor position 1.\n\nThe maximum L/D ratio for the wing-nacelle arrangement with external radiator is 16.6, or about 15 percent lower than for the bare-wing model. Similar data were not obtained for this model without radiators.\n\nPitching moments.- The power-off pitching-moment coefficients and the static longitudinal-stability characteristics of the model do not vary widely for all the arrangements tested. The slopes of the pitching-moment curves for the bare-wing and the enclosed-engine arrangements are slightly higher than those for the wing-nacelle model.\n\nPROPELLIVE AND OVER-ALL EFFICIENCIES\n\nEngine-propeller combinations should be compared by means of an over-all efficiency factor including both drag and propulsive efficiency. In this report the over-all efficiency is defined as the ratio of the power that would be required for the bare-wing model at a given speed, to the power input actually required at this speed for the particular propeller-wing combination.\n\nThe over-all efficiency of the bare-wing model is therefore 100 percent and, for an engine-propeller combination, is given by\n\n$$\n\\eta_T = \\eta \\left( \\frac{C_{D_W}}{C_{D_C}} \\right)\n$$", "timestamp": "2026-07-19T18:20:12.957008+00:00"}
{"citation_id": "19930091714", "source_url": "https://ntrs.nasa.gov/api/citations/19930091714/downloads/19930091714.pdf", "page_number": 32, "total_pages": 36, "image_filename": "19930091714_p32.jpg", "text": "28\nREPORT NO. 639—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nTABLE I (a)\nEXAMPLE 1, CONTROLLABLE PROPELLER\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 | 13 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| $\\frac{V}{nD}$ | $\\beta_1$ for $C_{P_1}=0.0582$ (deg.) | $C_{T_1}$ for $\\beta_1$ | $\\frac{C_{T_1}}{C_{T_1} \\text{ (at stall)}}$ | $\\frac{C_P}{C_P \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | $C_{T_2}$ | $\\beta_2$ for $C_{P_2}$ (deg.) | $C_{T_2}$ for $\\beta_2$ | $\\frac{C_T}{C_T \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | Thrust (lb.) | Air speed (m.p.h.) | Remarks |\n| 0 | 19.9 | 0.0870 | 1.01 | 1.165 | 0.0465 | 18.5 | 0.0860 | 0.995 | 0.0855 | 2.100 | 0 | Blades stalled. |\n| .1 | 20.2 | .0860 | 1.00 | 1.165 | .0465 | 18.9 | .0830 | .995 | .0845 | 2.075 | 19.3 | |\n| .2 | 20.7 | .0860 | 1.00 | 1.165 | .0465 | 19.0 | .0845 | .995 | .0840 | 2.060 | 38.8 | |\n| .3 | 20.7 | .0855 | 1.00 | 1.165 | .0465 | 19.0 | .0820 | .995 | .0815 | 2.000 | 58.2 | |\n| .4 | 20.6 | .0822 | .96 | 1.160 | .0467 | 18.7 | .0760 | 1.000 | .0760 | 1.870 | 77.5 | |\n| .5 | 20.6 | .0755 | .88 | 1.150 | .0471 | 19.0 | .0695 | 1.010 | .0705 | 1.727 | 97.0 | |\n| .6 | 20.9 | .0680 | .79 | 1.120 | .0483 | 19.6 | .0625 | 1.035 | .0642 | 1.575 | 116.5 | Blades not stalled. |\n| .7 | 21.6 | .0620 | .73 | 1.120 | .0483 | 20.3 | .0555 | 1.060 | .0589 | 1.445 | 135.5 | |\n| .8 | 22.3 | .0550 | .64 | 1.130 | .0479 | 21.2 | .0490 | 1.120 | .0550 | 1.350 | 155.5 | |\n| .9 | 23.3 | .0480 | .56 | 1.130 | .0479 | 22.3 | .0435 | 1.130 | .0495 | 1.215 | 175.0 | |\n| 1.0 | 24.6 | .0450 | .53 | 1.130 | .0479 | 23.7 | .0365 | 1.130 | .0446 | 1.095 | 194.0 | |\n\nTABLE I (b)\nEXAMPLE 2, CONTROLLABLE PROPELLER, DESIGN A\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 | 13 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| $\\frac{V}{nD}$ | $\\beta_1$ for $C_{P_1}=0.1097$ (deg.) | $C_{T_1}$ for $\\beta_1$ | $\\frac{C_{T_1}}{C_{T_1} \\text{ (at stall)}}$ | $\\frac{C_P}{C_P \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | $C_{T_2}$ | $\\beta_2$ for $C_{P_2}$ (deg.) | $C_{T_2}$ for $\\beta_2$ | $\\frac{C_T}{C_T \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | Thrust (lb.) | Air speed (m.p.h.) | Remarks |\n| 0 | 22.5 | 0.1420 | | | | | | | | | 0 | Blades stalled. |\n| .1 | 22.5 | .1410 | | | | | | | | | 21.8 | |\n| .2 | 23.0 | .1410 | | | | | | | | | 43.6 | |\n| .3 | 23.3 | .1405 | 1.00 | 1.240 | 0.0884 | 20.5 | 0.1320 | 0.970 | 0.1280 | 2.003 | 65.3 | |\n| .4 | 23.5 | .1395 | 1.00 | 1.240 | .0884 | 20.8 | .1280 | .970 | .1240 | 1.945 | 87.1 | |\n| .5 | 24.0 | .1330 | .95 | 1.230 | .0892 | 21.3 | .1190 | 1.000 | .1190 | 1.855 | 108.8 | |\n| .6 | 24.5 | .1270 | .91 | 1.220 | .0900 | 22.2 | .1095 | 1.020 | .1115 | 1.750 | 130.5 | |\n| .7 | 25.3 | .1180 | .84 | 1.200 | .0914 | 23.0 | .1010 | 1.060 | .1070 | 1.678 | 152.5 | Blades not stalled. |\n| .8 | 26.1 | .1090 | .78 | 1.185 | .0925 | 24.3 | .0930 | 1.090 | .1012 | 1.588 | 174.0 | |\n| .9 | 27.3 | .1000 | .71 | 1.170 | .0937 | 25.5 | .0870 | 1.115 | .0972 | 1.522 | 196.0 | |\n| 1.0 | 28.0 | .0910 | .65 | 1.157 | .0948 | 27.0 | .0800 | 1.135 | .0909 | 1.425 | 218.0 | |\n| 1.1 | 30.3 | .0840 | .60 | 1.145 | .0958 | 28.2 | .0740 | 1.150 | .0852 | 1.335 | 239.0 | |\n| 1.2 | 30.5 | .0780 | .56 | 1.138 | .0965 | 29.6 | .0690 | 1.140 | .0786 | 1.231 | 261.0 | |\n| 1.3 | 32.0 | .0730 | .53 | 1.130 | .0970 | 31.0 | .0640 | 1.135 | .0727 | 1.140 | 283.0 | |\n\nTABLE I (c)\nEXAMPLE 3, CONTROLLABLE PROPELLER\n\n| 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 | 13 |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| $\\frac{V}{nD}$ | $\\beta_1$ for $C_{P_1}=0.1097$ (deg.) | $C_{T_1}$ for $\\beta_1$ | $\\frac{C_{T_1}}{C_{T_1} \\text{ (at stall)}}$ | $\\frac{C_P}{C_P \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | $C_{T_2}$ | $\\beta_2$ for $C_{P_2}$ (deg.) | $C_{T_2}$ for $\\beta_2$ | $\\frac{C_T}{C_T \\left(\\frac{V}{V_s}=0.5\\right)}$ | $\\frac{C_{T_1}}{C_{T_2}}$ | Thrust (lb.) | Air speed (m.p.h.) | Remarks |\n| 0 | 22.6 | 0.1820 | 1.00 | 1.305 | 0.084 | 19.0 | 0.1730 | 0.75 | 0.1298 | 2.035 | 0 | Blades stalled. |\n| .1 | 22.6 | .1810 | 1.00 | 1.305 | .084 | 19.1 | .1640 | .73 | .1230 | 1.928 | 21.8 | |\n| .2 | 22.8 | .1770 | .98 | 1.305 | .084 | 19.2 | .1530 | .70 | .1160 | 1.818 | 43.6 | |\n| .3 | 22.9 | .1680 | .94 | 1.305 | .084 | 19.5 | .1400 | .79 | .1105 | 1.720 | 65.3 | |\n| .4 | 23.4 | .1550 | .86 | 1.305 | .084 | 19.5 | .1290 | .88 | .1135 | 1.780 | 87.1 | |\n| .5 | 23.3 | .1400 | .79 | 1.305 | .084 | 20.1 | .1170 | .98 | .1133 | 1.780 | 108.8 | |\n| .6 | 23.6 | .1300 | .72 | 1.330 | .0836 | 21.0 | .1060 | 1.05 | .1105 | 1.730 | 130.5 | Blades not stalled. |\n| .7 | 24.6 | .1190 | .66 | 1.355 | .0819 | 21.6 | .0960 | 1.13 | .1059 | 1.660 | 152.5 | |\n| .8 | 25.4 | .1090 | .61 | 1.360 | .0807 | 22.6 | .0820 | 1.20 | .0984 | 1.540 | 174.0 | |\n| .9 | 26.4 | .0990 | .55 | 1.350 | .0813 | 24.1 | .0700 | 1.22 | .0852 | 1.460 | 196.0 | |\n| 1.0 | 27.6 | .0910 | .51 | 1.330 | .0825 | 25.2 | .0700 | 1.25 | .0862 | 1.350 | 218.0 | |\n| 1.1 | 28.8 | .0840 | .47 | 1.310 | .0838 | 27.0 | .0640 | 1.22 | .0791 | 1.240 | 239.0 | |\n| 1.2 | 30.3 | .0770 | .43 | 1.290 | .0850 | 28.3 | .0590 | 1.23 | .0729 | 1.145 | 261.0 | |\n| 1.3 | 31.9 | .0710 | .40 | 1.275 | .0861 | 30.0 | .0550 | 1.24 | .0682 | 1.070 | 283.0 | |", "timestamp": "2026-07-19T18:20:23.350320+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 28, "total_pages": 43, "image_filename": "19930094538_p28.jpg", "text": "26 N.A.C.A. Technical Memorandum No. 878\n\nTABLE IV\n\nComparison of Theoretical and Experimental Failing Torque Values\n\n| Cylinder No. | Experimental failing torque $T_B$ kg cm | Theoretical failing torque and discrepancies | | | | | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| | | **Complete tension field** | | **Incomplete tension field** | | | |\n| | | Theoretical failing moment kg cm | Discrepancy percent | Free value $R_0$ | $\\delta_g$ | Theoretical failing moment kg cm | Discrepancy percent |\n| II | 198,500 | - | - | - | - | - | - |\n| III | 307,000 | 309,800 | +1.0 | 0.85 | 43 | 318,200 | + 3.8 |\n| IV | 231,000 | 198,200 | -16.2 | .72 | 60 | 250,000 | + 8.1 |\n\nIV. INVESTIGATION OF THE FAILING TORQUE\n\n1. The Failing Torque and Its Solution\n\nThe failing torque is, according to table IV, highest on cylinder No. III; lowest on cylinder No. II. In all cases the failure takes place as column effect of the stringers. On cylinder No. II (figs. 22 and 23), the stringers are twisted, as a result of which the effective moment of inertia of the sections continues to drop as the load increases. Cylinder No. III (fig. 24), being of closed profile form, does not manifest this phenomenon; failure takes place as the result of denting of the section flanges. The webs of the open-hat sections of cylinder No. IV (fig. 25), twist individually, while the back of the section is dented in. The strips of the sheets riveted to the sections are dented on the very points of failure as a result of the buckling of the sections; on the remaining points they remain smooth. This fact permits the application of the assumption of effective sheet to the solution of the ultimate torque.\n\nIt is logical to compute the ultimate torque on the basis of buckling-bending tests with shell panels. In the following the bulkheads are assumed to be strong enough to effect a buckling of the stringers in one wave each in every bulkhead panel.", "timestamp": "2026-07-19T18:20:26.515255+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 21, "total_pages": 32, "image_filename": "19930094552_p21.jpg", "text": "N.A.C.A. Technical Memorandum No. 864 19\n\ntially on the flexural strength of the \"main spars.\" In the arching type of loading with the two forward bulkheads riveted, the main spars transmit the force to the forward bulkhead in a radial direction. The support of the main spars on the flexurally weak bulkheads at the free end leads to a strong elliptical bulging of the latter. The same effect occurs as a result of the radial loading of the spars through the peripheral stress components if the forward bulkheads are not attached to the skin.\n\nRiveting of the two forward bulkheads to the skin produces a somewhat more rapid rate of decrease of the forces in the main spars in the arching loading condition and the setting up of breaks in the curves of longitudinal force $P_x(x)$. The more rapid rate of decrease is to be explained by the fact that the skin contributes an effective supporting width when the bulkheads are under a bending load. The resistance of the shell against arching is not increased to any considerable extent by riveting since, with a stiff ring mounted at the end, the forces transferred from the main spars through the riveting are not changed by the riveting of the forward bulkheads. The end ring has the effect of strongly relieving the load from the bulkheads at the free end.\n\nThe buckling of the sheet had no effect in reducing the forces in the main spars within the range investigated (in bending, up to 2.5 times the buckling load and in arching, up to 1.8 times the buckling load).\n\nThe results of the tests suggest the necessity for an extension of the simple shear field by the addition of the transverse stresses and indicate that the bending strength of the stringers may be of considerable effect on the loading of the bulkheads.\n\nThe tests should be further extended to the case - more difficult to compute - where the cutaway of the shell lies only on one side and the shell at the side lying opposite is of framework construction.\n\nTranslation by S. Reiss,\nNational Advisory Committee\nfor Aeronautics.", "timestamp": "2026-07-19T18:20:26.706294+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 16, "total_pages": 24, "image_filename": "19930091716_p16.jpg", "text": "```markdown\n12\nREPORT NO. 641—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n### FRICTION TORQUE OF ENGINES\n\nRelation to problems.—In many cases the friction of the engine has a large effect on the values of thrust coefficient under which the propeller operates. It is then obvious that, for the ready use of the data herein presented, some information regarding the friction torque of engines is required. As particular and exact engine-friction data are seldom available for the solution of negative-thrust problems, it is considered necessary to include in this report sufficient data to permit an intelligent estimate of friction torque to be made.\n\nEngine-friction data.—Figure 14 shows friction-torque curves, obtained from various sources, for eight modern aircraft engines covering a fairly wide range of power and displacement. Some of these curves have been extrapolated to bring them to 1,000 r. p. m. From these curves may be judged the quality of the assumption that the friction torque is a linear function of $N_e$, which can be represented by the relation $Q_f/N_e=K$ for any particular engine. This equation indicates that the torque for any engine may be represented by a straight line through the origin, as the one drawn in for engine 3 in figure 14.\n\n<!-- Image (209, 206, 811, 661) -->\n\nFIGURE 14.—Friction-torque curves for eight typical airplane engines.\n\n| Engine | $\\Delta$ | hp. | $N_e$ | $Q_f/N_e$ |\n| :--- | :--- | :--- | :--- | :--- |\n| 1 | 1,823 | 768 | 1,950 | -0.120 |\n| 2 | 1,820 | 700 | 1,950 | -.125 |\n| 3 | 1,535 | 850 | 2,150 | -.085 |\n| 4 | 1,535 | 750 | 2,500 | -.065 |\n| 5 | 1,340 | 450 | 2,100 | -.075 |\n| 6 | 985 | 300 | 2,000 | -.050 |\n| 7 | 975 | 330 | 2,000 | -.050 |\n| 8 | 499 | 145 | 2,050 | -.025 |\n\nThe curves in figure 14 indicate that this approximation is not far wrong for values of $N_e$ from 1,000 to 2,200 and for the temperature and pressure conditions under which the tests were run. It is certain, however, that this approximation does not hold at low values of $N_e$ for it assumes that the torque becomes zero at $N_e=0$. The torque does not become zero at $N_e=0$, as is indicated by the extrapolated curve for engine 4,\n```", "timestamp": "2026-07-19T18:20:43.461290+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 3, "total_pages": 22, "image_filename": "19930091655_p3.jpg", "text": "REPORT No. 580\n\nHEAT TRANSFER TO FUEL SPRAYS INJECTED\nINTO HEATED GASES\n\nBy ROBERT F. SELDEN and ROBERT C. SPENCER\nLangley Memorial Aeronautical Laboratory\n\n103475—37——1", "timestamp": "2026-07-19T18:20:56.084155+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 5, "total_pages": 16, "image_filename": "19930094559_p5.jpg", "text": "N.A.C.A. Technical Memorandum No. 857 3\n\nwithin the test range, the pressure rise had to be taken into account. The tests showed, however, that the effect of the air velocity was by far predominant over that of the air pressure on the combustion process.\n\nFigures 3 to 8 show some of the flame photographs obtained. The fuel quantity of about 50 mg corresponds approximately to half-load of the engine that was used for comparison. The pictures are arranged below each other, in order of start of injection, referred to crankshaft angle of compressor ($360^\\circ$ = top dead center of compressor). The counterclockwise rotation of the fuel spray, corresponding to the air movement, is clearly visible. The rotational motion of the gas masses is particularly well brought out in figures 5 and 6. This motion can be made out also on the other pictures if the transport of the fuel by the air movement is compared from one picture to the next, or the pictures compared with those of figure 8, showing \"injections into still air.\"\n\nThe projection of the flame out through the connecting passage into the main combustion chamber of the cylinder, depends not only on the spray penetration but essentially on the direction of pressure drop between turbulence and main chambers. Due to the greater air capacity of the cylinder-shaped model chamber as compared with that of the engine, any considerable amount of projection of the flame is observable only with the greater fuel quantities. Compare figures 5 and 7 corresponding, respectively, to fuel quantities of 50 and 100 mg. The more rapid the rate of decrease of the flow into the chamber without combustion — i.e., the later the injection occurs after the maximum inflow velocity, the greater the flame projection. Since in these model tests the expansion process could not be simulated, it is to be expected that in the case of the engine and particularly also on account of the smaller volume and hence steeper pressure rise in the turbulence chamber, there will be a greater amount of projection of the flame from the turbulence chamber than is shown in these photographs.\n\nFigure 9 gives the numerical results obtained as a function of the start of injection. The ignition lag $z_v$ decreases with later start of injection corresponding to the pressure (or density) and velocity increase of the air, as shown in figure 2. To the ignition lags were added the times of complete combustion $z_{Br}$. For two different fuel quantities, two curves were obtained which", "timestamp": "2026-07-19T18:21:05.858996+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 12, "total_pages": 47, "image_filename": "19930093641_p12.jpg", "text": "10\n\nThe effective thrust of the propeller-body combination,  \n$T - \\Delta D$, is obtained from the measured data by means of  \nthe relation  \n\n$$\nR = D_c + \\Delta D - T\n$$\n\nFor tests without a wing behind the propeller, $T - \\Delta D$  \nis obtained from measurements of $D_c$ and $R$ for the same  \nangle of attack and dynamic pressure. When the propeller  \nis operated in front of or behind a wing, there are changes  \nin the lift as well as in the drag and jet-boundary corrections  \nthat should be credited to or charged against the propeller. The change in lift has been allowed for in these  \nresults by determining $D_c$ and $R$ at the same lift coefficient instead of at the same angle of attack. Since  \nhigher lift coefficients are reached with power on than  \noff, this method fails in the region of maximum lift; however, it is valid over the remainder of the useful flight  \nrange.\n\nPropulsive efficiencies are given for two lift coefficients of the model $C_L = 0.25$ and $0.70$, which corresponds approximately to the lift coefficients for high  \nspeed and climb. Of particular interest are the curves of  \nfigure 26, comparing the efficiencies for the five principal engine-propeller combinations. A blade angle of  \n$18\\text{-}1/2^\\circ$ was used for the comparative tests inasmuch as it  \nrepresents approximately the setting required to absorb  \nthe available power in the climb condition. At the high-speed lift coefficient (fig. 26), the maximum propulsive  \nefficiencies show a dispersion of only about 2 percent between all the combinations tested; the highest value,  \nnearly 80 percent, is given by the pusher and the lowest  \nvalue, 78 percent, by the conventional wing-nacelle arrangement. In sharp contrast are the values shown in figure 27 for the climb lift coefficient, in which there is a  \ndifference of 8 percent between the highest maximum efficiency, 83 percent for the enclosed-engine tractor position 1, and the lowest maximum efficiency, 75 percent for  \nthe conventional wing-nacelle arrangement. The pusher,  \ntractor position 2, and tractor position 3 follow tractor  \nposition 1 in decreasing order of merit.\n\nThe effect of blade angle for the conventional wing-nacelle arrangement is shown in figure 28. With increasing blade angle, the propulsive efficiency increases up to  \n$\\beta = 23\\text{-}1/2^\\circ$ for the high-speed lift coefficient and re-", "timestamp": "2026-07-19T18:21:09.600970+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 8, "total_pages": 17, "image_filename": "19930094535_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 881\n\nD. Polyacrylic acid esters\n\nPlates of acrylic acid esters are formed by casting into various shapes and in thicknesses of 0.5 to 10 mm.\n\nPrincipal Properties of Polyacrylic acid esters\n\nSpecific weight . . . . . . 1.18\n\nTensile strength . . . . . . 750 kg/cm²\n\nBending strength, about . . . 1,100 kg/cm²\n\nElasticity modulus . . . . . 28,000 kg/cm²\n\nImpact bending strength . . . 15-30 cm kg/cm²\n\nHeat resistance according to Martens, about . . . . . . 50° C.\n\nExpansion coefficient . . 130 x 10⁻⁶\n\nWORKING OF THE MATERIALS\n\nThe four artificial materials indicated above, namely, cellulose nitrate, cellulose acetate, polymer mixtures, and polyacrylic acid esters may be worked in almost the same manner. In the case of celluloid, however, particular caution must be applied on account of its inflammability.\n\nThese materials can readily be shaped at low heat without introducing any stresses. The heating can be done by warm metal plates, hot air, or hot water - a temperature of 75° to 100° C. being required. The heated material is drawn over heated forms of sheet metal, wood, or glass and must be cooled in the form. The material can easily be bored, mechanical means being best suited for the purpose. Care must be taken in order to avoid the formation of scratches, soft layers, such as paper, being used under the materials. Turning is possible on any lathe. The materials can be cut along straight or slightly curved lines with a small band or fret saw, or in the same manner as glass, except that instead of the expensive diamond, a fine steel point may be used. The steel point must be applied with some pressure, and for thicknesses above 1.3 the scratching", "timestamp": "2026-07-19T18:21:16.720861+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 13, "total_pages": 21, "image_filename": "19930094534_p13.jpg", "text": "12 N.A.C.A. Technical Memorandum No. 892\n\nquickest results. And the very first experimental machine developed on this principle proved successful. The necessary shaping tools themselves remained as before. The drive gives the upper tool holder a stroke of 1,2 mm with a frequency of 700 blows per minute. The action of the individual pressures is controlled by the spacing of the two tools at dead center and regulated by vertical adjustment of the bottom tool. This adjustment is effected by a simple screw spindle with foot-pawl operation. The adjusting force necessary for this is moderate, since adjustment can be made only in the pressure-free part of the stroke and the end pressures are absorbed by self-locking of the spindle.\n\nThe upper tool being movable upward and sideways, the drawing pressure can be applied at any desired point of the section. This adjustment is obtained by supporting the upper tool holder in a double eccentric whose two adjusting levers are easily operable. Since with a stroke of only 1,2 mm the clearance in the joints would permit of no satisfactory operation and cause disturbing noises, the upper tool holder is provided with a spring of high oscillation frequency, which acts against the driving pressure.\n\nThe framework is designed for 50-ton maximum pressure and fitted with a safety device for 30-ton ultimate load. The motive power for this range is 5 horsepower. The machine not only meets the requirements as regards absence of noise but actually increases the quality of production considerably. Jerky feeding by hand is impossible because of the high frequency; the section passes through almost automatically if the section is pushed ahead steadily, affording a smooth, uniform surface. Judged by the experiences gained thus far, a 50-percent greater output over the conventional air hammers is anticipated. Not being bound to the uneconomical compressed air, the machine can also be used for portable plants; it likewise eliminates keeping a number of rarely used form-fitting tools in stock, which is of vital importance in experimental design. For instance, U-shaped arcs of 2 mm sheet thickness, with an 80 mm web height and 30 mm flange width heretofore shaped from sheet over form-fitting tools can be obtained more economically by driving from previously beveled channel sections.", "timestamp": "2026-07-19T18:21:17.009680+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 6, "total_pages": 18, "image_filename": "19930094551_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 865 5\n\nand employing the cited metal tubing, it was possible to lower these transverse oscillations to a minimum. This arrangement damps out even oscillations set up in gusty weather remarkably quickly without causing any appreciable disturbance of the head itself.\n\nHaving succeeded in the matter of weight of pressure head and type of tubing to insure in principle perfect performance up to speeds of almost 400 km/h the position of the pressure head with respect to the airplane was determined at speeds of from 210 km/h to 350 km/h with tubing of 12, 15, 20, and 24 m length.\n\nWith the notation of figure 1, the depth and trail of the pressure head follow from the relations:\n\n$$\nz = a \\frac{\\sin k_1 \\sin k_2}{\\sin (k_1 - k_2)} \\text{ [m]} \\tag{1}\n$$\n\n$$\nx = a \\frac{\\cos k_1 \\sin k_2}{\\sin (k_1 - k_2)} \\text{ [m]} \\tag{2}\n$$\n\nFigure 3 shows the depth z from the measuring base a parallel to the longitudinal airplane axis plotted against the speed, with tubing length as parameter. With tubing of 20 m length, the normal depth 9.7 m drops to 6.5 m when the speed changes from 210 to 350 km/h, which, extrapolated to 400 km/h, still leaves a depth of 6.0 m. This length should undoubtedly suffice to reduce the error in reading resulting from the falsification of the static pressure to a practically insignificant magnitude (reference 1).\n\nThe effect of the suspension length on the normal distance from the reference base is comparatively small within the explored limits according to figure 3. Doubling the length from 12 to 24 m increases, at 300 km/h, the distance from 5.5 to 7 m, i.e., a 100-percent-length change produces only a 27 percent change in depth. With greater hose lengths the change becomes even less, because the angle of exit of the hose from the airplane becomes consistently less on account of the greater total drag.\n\nMuch more important is the influence of the suspension length upon the position of the head behind the airplane.", "timestamp": "2026-07-19T18:21:18.974646+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 22, "total_pages": 32, "image_filename": "19930094552_p22.jpg", "text": "20 N.A.C.A. Technical Memorandum No. 864\n\nREFERENCES\n\n1. Ebner, H., and Köller, H.: Über die Einleitung von Längskräften in versteifte Zylinderschalen. Jahrbuch 1937 der deutschen Luftfahrtforschung. R. Oldenbourg 1937, München-Berlin.\n\nEbner, H., and Köller, H.: Calculation of the Load-Distribution in Stiffened Cylindrical Shells. T.M. No. 866, N.A.C.A., 1938.\n\n2. Wagner, H., and Simon, H.: Über die Krafteinleitung in dünnwandige Zylinderschalen. Luftfahrtforschung, vol. 13, no. 9, 1936, pp. 293-308.\n\n3. Ebner, H.: The Strength of Shell Bodies - Theory and Practice. T.M. No. 838, N.A.C.A., 1937.\n\n4. Redshaw, C.: The Elastic Instability of a Thin Curved Panel Subjected to an Axial Thrust, Its Axial and Circumferential Edges Being Simply Supported. R.& M. No. 1535, British A.R.C., 1934.", "timestamp": "2026-07-19T18:21:19.153232+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 29, "total_pages": 43, "image_filename": "19930094538_p29.jpg", "text": "N.A.C.A. Technical Memorandum No. 878 27\n\nThe compression in a stringer is $P_L = \\bar{\\sigma}_x F_L$; the static cross load on it between two bulkheads is $Q = p t$. The ratio $\\lambda = P_L/Q$ between both follows from equations (12b) and (20) as:\n\n$$\n\\lambda = - \\frac{b_1}{\\phi t} R \\frac{\\delta \\cot \\alpha - 1}{\\delta \\tan \\alpha - 1} \\quad \\text{with} \\quad \\delta = \\frac{T}{T_0} \\tag{30}\n$$\n\nShells with thin skin and strong stiffeners may be expressed with $\\lambda \\sim - \\frac{b_1}{\\phi t} R \\cot^2 \\alpha$ when approaching failure. The ratio $\\lambda$ therefore depends upon $\\delta$; this function $\\lambda(\\delta)$ can be computed if $\\alpha(\\delta)$ and $R(\\delta)$ are known. The axial load on each stiffener is:\n\n$$\nP_L = - \\delta R F_L T_0 (\\delta \\cot \\alpha - 1) \\tag{31}\n$$\n\nIn the buckling-bending tests the shell panels, being of a length equal to the bulkhead spacing, are subjected to a uniformly distributed cross load. The ends of the sections should be built in. Figure 26 indicates the loading and the experimental arrangement. The ratio $\\lambda = P_L/Q$ is varied during the tests; the failing loads $P_{LB}$ are plotted against $\\lambda$, figure 27. Given the relation $P_{LB}(\\lambda)$ and the function $\\lambda(\\delta)$ (fig. 28), a function $P_{LB}(\\delta)$ is obtainable that gives the axial loads still supported under the different stress conditions $\\delta$ on the basis of the related load conditions $\\lambda$. The intersection of the curves $P_L(\\delta)$ and $P_{LB}(\\delta)$ defines the stress condition $\\delta_B$, at which the shell should fail as a result of exhausted carrying capacity of the longitudinal sections. The ultimate torque is $T_B = \\delta_B T_0$. The argument holds only for the case that the axial load $P_L$ remains below the Euler buckling load for supported bar ends.\n\n2. The Buckling-Bending Tests on Shell Panels\n\nThe shell panels consisted of two longitudinal sections bordering a riveted skin panel, and resembled a sheet panel of the respective cylinder. The length of the panels equaled the bulkhead spacing of the cylinder. The skin section was not fastened to the clamping angles and was not subject to any outside compression. The cross load (fig. 26) was distributed as evenly as possible over the length by means of wooden strips and felt liners.", "timestamp": "2026-07-19T18:21:21.108374+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 19, "total_pages": 43, "image_filename": "19930094544_p19.jpg", "text": "N.A.C.A. Technical Memorandum No. 872 17\n\ndifferent types. In the first type the intersecting booms are riveted directly together. With this type one recognizes that, as a result of the eccentric attachments of the individual members and of the stiff construction of the joint, stress concentrations occur, which, however, are in general of no great disadvantage, since they occur only locally. The other type seeks to reduce these secondary stresses, since as much as possible it brings the members together at one point in special junction members. This type has the advantage in assembling, that all members can be completed in their correct lengths and then screwed up. The structural design is, however, more difficult and also involves more weight.\n\nBecause of these considerations German airship building has thus far not departed from the stiff riveting of the joints. Figure 43 shows a typical joint, as it occurs in the construction of the LZ 127. The longitudinal girder with the downward pointing apex passes through the ring girder. Underneath the attachment plate for field assembly is visible. Also the girders of the \"Akron\" are riveted at the joints. Figure 44 shows an inner joint of the main ring. Here especially simple attachments result from the rectangular design of the girders.\n\nIn the construction of the R 100, special joint members (fig. 46) have been riveted together, on to which the boom tubes of the longitudinal and ring girders are screwed by means of sleeve nuts. Such a joint completed is seen in figure 45, which again shows the continuity of a longitudinal girder at the ring corner.\n\nA ring joint of the R 101 looks entirely different (fig. 47). The boom tubes of the ring struts are brought together in pyramid form and end in a light metal casting (fig. 48), which is held by the fork-like ends of the tubes of the inner ring booms. Also the wire attachments in the R 101 are worked out in an unusual manner. The wires are poured into sleeves, which are screwed into casings. The casings are swivel-fastened to a steel plate, which can turn around a bolt set in the joint casting.\n\nIn the \"Graf Zeppelin\" as well as in the \"Akron\" the ends of wires are looped, served with small wire and then soldered. The new structure of the LZ 129 has departed from this type of wire terminal for the bracing of the main rings. The wires, which here in places go to wire diameters up to 8 mm, end in so-called \"Heddernheimer\" casings,", "timestamp": "2026-07-19T18:21:22.854929+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 41, "total_pages": 50, "image_filename": "19930094543_p41.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\n\nFigs. 1,3,4\n\nConstantan\nIron\n\n[Figure: Cross-section of a thermocouple junction showing Constantan and Iron wires embedded in a metal block.]\n\nFigure 1\n\nHot spot\n\n[Figure: Schematic diagram of an electrical circuit with a transformer, resistors, and a hot spot indicator.]\n\n110V∞\n\nFigure 3\n\nG\n\nE\n\n[Figure: Schematic diagram of a galvanometer (G) connected to a circuit with a switch and a rotating element.]\n\nFigure 4", "timestamp": "2026-07-19T18:21:27.385466+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 23, "total_pages": 32, "image_filename": "19930094552_p23.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFigs. 1,2\n\nSections through test cylinder.\n\nSection C-D\nSection A-B\nMain spars\n\nLongitudinal section\nMeasuring sections\n\nLongitudinal stiffener sections\n\n3; 7; 12; 16 Measuring stations 5; 14\nRivet diameter 4 mm\n\n1; 2; 4; 6; 8; 9; 10; 11;\n13; 15; 17; 18\nMeasuring stations\nRivet diameter 3mm\n\nBulkhead ring sections\n\nMeasuring station e; f.\na; b; c; d; g\nMeasuring stations\n\nMeasuring station\n\nFigure 1.- Test cylinder with stiffener sections.\n\nBending forces\nArching forces\n\nFigure 2.- Loading conditions.", "timestamp": "2026-07-19T18:21:48.547404+00:00"}
{"citation_id": "19930094557", "source_url": "https://ntrs.nasa.gov/api/citations/19930094557/downloads/19930094557.pdf", "page_number": 18, "total_pages": 20, "image_filename": "19930094557_p18.jpg", "text": "[No readable text detected]", "timestamp": "2026-07-19T18:21:48.924061+00:00"}
{"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 16, "total_pages": 20, "image_filename": "19930091707_p16.jpg", "text": "12\nREPORT NO. 632—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n<!-- Image (128, 79, 907, 543) -->\n\nFIGURE 15.—Diagram of $\\sigma_{cr}$, $1/\\delta$ for duralumin (transverse loading).\n\npoints because the stress-strain curves of the material they represent are not even approximately affinely related to the stress-strain curves of the other material.\n\nIt will be noted that smooth curves drawn to represent the $\\sigma$, $1/\\delta$-data would rise concave upward to meet the points for which $1/\\delta$ is greater than 19. These points represent the strengths of relatively thick specimens, and it is to be expected that the strengths of such specimens would increase rapidly with decrease in ratio of diameter to thickness. If $d/t=2$, the tube becomes a solid bar, the crinkling strength becomes infinite, and the bending strength becomes very high, depending now on the tensile strength of the material. These high values of crinkling strength and bending strength, however, have no practical significance since the deformations required to obtain them are so great as to be quite intolerable in a structure. The more or less abrupt rise in strength at low values of the ratio of diameter to thickness is analogous to that which occurs in columns at low values of the ratio of slenderness.\n\nEmpirical curves have been fitted to the $\\sigma$, $1/\\delta$-data of figures 9, 11, 13, and 15. In doing so, the points for which $1/\\delta$ was greater than 19 have not been used, in accordance with the preceding discussion, nor have the three sets of points for the chromium-molybdenum-steel specimens previously discussed been taken into account, since they represent essentially a different material. The omission from consideration of all these points is on the safe side. All four curves are hyperbolas. They are shown solid in the figures for the range covered by the tests, and they are extended as dotted curves. The crinkling strength, in nondimensional form, of the chromium-molybdenum-steel tubes was found to be given by\n\n$$\n\\sigma_{cr\\theta} = \\frac{1.315 \\frac{1}{\\delta_y}}{1 + 1.98 \\frac{1}{\\delta_y}}, \\quad \\frac{1}{\\delta_y} > 5 \\quad (6)\n$$\n\nThe crinkling strength, in nondimensional form, of the duralumin tubes was found to be given by\n\n$$\n\\sigma_{cr\\theta} = 1.631 - \\frac{3.35}{\\frac{1}{\\delta_y} + 2.75}, \\quad \\frac{1}{\\delta_y} > 2.5 \\quad (7)\n$$", "timestamp": "2026-07-19T18:21:55.716500+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 30, "total_pages": 43, "image_filename": "19930094538_p30.jpg", "text": "28 N.A.C.A. Technical Memorandum No. 878\n\nThe test procedure and the aspect of the break were the same as for the cylinder test. The sections of the panels belonging to cylinder No. III, failed after indentation of the flanges; but those of panel 13a (cylinder IV) failed through twisting failure of the webs.\n\nThe tests indicated (fig. 27) on the panels with closed sections (of cylinder III), first a rapid, and subsequently a slower, increase in the falling load $F_{LB}$ with the load ratio $\\lambda = P_T/Q$. For the panels with open sections (cylinder IV), the ultimate axial load within the explored range ($\\lambda$ between 3 and 4) is very little affected by $\\lambda$. All experiments indicate a wide scatter zone. For the solution of the twisting stiffness the lower margin of this zone is decisive, since the failure is always initiated by the weakest of a large number of identically stressed sheet panels.\n\nLoading two individual stiffeners instead of a panel, the former - if of closed section - proved stronger than the sheet panels in spite of the absence of the skin section, because with the latter the buckling of the section flanges is initiated by the skin. On the panels of open sections the skin had no effect on the strength.\n\n3. Checking the Calculation against the Test Data\n\nThe calculation of the breaking torque included both the complete and the incomplete tension field. The data appended in table IV allows a comparison with the experimental results.\n\nPremised on a complete tension field, the breaking torque of cylinder No. III was computed exact at 1.5 percent, while on cylinder No. IV, it was computed 16.2 percent too low. From this it follows that with thin skin and strong stiffeners the assumption of incomplete tension field leads to a sufficiently exact solution of the strength. For somewhat thicker skin and weaker stiffeners, the values by this calculation method are too low, since the tension field even at failure is still not completely stretched.\n\nThe calculating method for incomplete tension field contains the free values $R_0$ and $\\theta_z$ which on cylinder No. IV could be so chosen that the theoretical and experi-", "timestamp": "2026-07-19T18:21:57.977569+00:00"}
{"citation_id": "19930094544", "source_url": "https://ntrs.nasa.gov/api/citations/19930094544/downloads/19930094544.pdf", "page_number": 20, "total_pages": 43, "image_filename": "19930094544_p20.jpg", "text": "18 N.A.C.A. Technical Memorandum No. 872\n\nwhich are turned up over the wire ends on the wire-drawing frame. According to tests which were conducted at the DVL, these casings represent an exceptional terminal joint (26). In the rings of the LZ 129, now under construction, an especially interesting attachment of the wire bracing to the ring corners has been developed (fig. 49). It has for its object the leading of the wire forces as centrally as possible into the ring joints, in order to reduce torsion and lateral bending stresses in the ring girders. The wires coming into the joint are brought together on a steel member, the so-called \"spreader.\" Around this is laid an endless cable strop, which is led over a formed part, the so-called \"whip.\" This formed part swings on a bolt, which is placed at the junction point of the ring and longitudinal girders.\n\n3. Materials.\n\nIn the structures of LZ 127, LZ 129, \"Akron,\" and R 100, duralumin is used as structural material. In the R 101 a mixed construction has been adopted, in which the boom tubes of the longitudinal girders are worked out in steel. The question, which of the two materials mentioned is more suitable for the airship frame is difficult to decide theoretically. If one compares the pure efficiency factors for columns, then, to be sure, duralumin shows up the better; one should not forget, however, that in view of the compact design and the possibility of welding in the case of steel construction the joints turn out lighter. With the size of present-day airship structures we have undoubtedly come into a range where steel, especially in the form of weldable tubes, comes into the picture as a serious competitor of duralumin, which is preferably used in open sections on account of riveted attachments.\n\nIn table 1* are assembled the duralumin alloys heretofore used in airship structures. Hardness 1 signifies: cold rolled after refining. The corresponding values can also be applied to drawn sections, since approximately the same strengthening results from drawing. The first series\n\n*The table is taken from the paper by Dr. Ing. Brenner: \"Die Auswirkung neuerer Erkenntnisse der Werkstofforschung auf den Luftfahrzeugbau\" (\"The Development of New Conceptions of Material Research in Aircraft Construction\"), appearing in the DVL-Jahrbuch 1933.", "timestamp": "2026-07-19T18:22:02.712199+00:00"}
{"citation_id": "19930094535", "source_url": "https://ntrs.nasa.gov/api/citations/19930094535/downloads/19930094535.pdf", "page_number": 9, "total_pages": 17, "image_filename": "19930094535_p9.jpg", "text": "8 N.A.C.A. Technical Memorandum No. 881\n\nmust be repeated several times. For breaking narrow pieces flat-nosed tongs are used and wider specimens may be broken along the groove by hand. The rough edge on the scratch side of the broken piece is removed by drawing off with a file or with the blade of the scratcher as otherwise small grains may fall off from the rough edge and cause abrasions. The polyacrylic acid esters may be glued to wood, metal, textiles, and rubber with the aid of special putties. Care must be taken that those parts which are not to be glued do not come in contact with the solutions and glues, since they may attack the material and injure the surface. Machines used in grinding glass are also suitable for these artificial materials. The grinding is best done on emery disks and high-speed emery bands of medium grain. Polishing is done on felt disks under slight pressure with some pumice-stone powder and much water added. Strong heating is to be avoided. If the material is to be stamped, forms similar to those used for leather are used. Long heating or temperatures above 80° C. are to be avoided.\n\nThe small surface hardness that is always found with the artificial transparent materials, may give rise to scratches and abrasions at very sandy or dusty flying fields. By polishing, however, these scratches may easily be removed. Figure 6 shows a pane of polyacrylic acid ester which was set into the frame of a test airplane and exposed for 100 flying hours to the action of the most varied weathering conditions. No scratches or abrasions of any importance are observed, and the light transmission was in no way impaired.\n\nA disadvantage is the very large expansion coefficient which in the case of the polyacrylic acid esters is 500 percent as great as that of aluminum; i.e., at 1-meter length the material will shrink, for a lowering of temperature by 40° C., about 5 millimeters whereas aluminum will change by only 1 millimeter. In mounting these artificial glass panes, care should therefore be taken that they are not rigidly fixed. In riveting, the orifices are to be bored 3 to 4 millimeters larger, making use of cover strips or washers. A very suitable method is framing in U-shaped edge strips. Further methods of attachment are illustrated in figure 7. In order to assure a uniform support, an intermediate layer of leather or rubber is, in all cases, advisable.\n\nWhile the polyacrylic acid esters, from the point of", "timestamp": "2026-07-19T18:22:04.479717+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 42, "total_pages": 50, "image_filename": "19930094543_p42.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\nFigs. 2,5,26,27\n\n[Figure: Diagram of a mechanical apparatus with labeled parts A, B, C, D, E, F]\nFigure 2\n\n[Figure: Schematic diagram of an electrical circuit with a galvanometer (G), a hot spot, and a thermometer in a beaker]\nThermometer\nHot spot\nFigure 5\n\n[Graph: Pressure in kg/cm² vs. Volumes in liters, showing a curve starting at approximately 30 kg/cm² and decreasing to near zero as volume increases from 0 to 1 liter]\nFigure 26.- Diagram No. 1 transformed in p.v. axes.\n\n[Graph: Pressure in kg/cm² vs. Volumes in liters, showing two curves starting at approximately 30 kg/cm² and decreasing, with one curve dipping below the other before both approach zero as volume increases from 0 to 1 liter]\nFigure 27.- Diagram No. 4 transformed in p.v. axes for studying the variations in polytropic coefficient during compression and expansion.", "timestamp": "2026-07-19T18:22:05.416672+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 4, "total_pages": 22, "image_filename": "19930091655_p4.jpg", "text": "```markdown\n# NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\nHEADQUARTERS, NAVY BUILDING, WASHINGTON, D. C.\n\nLABORATORIES, LANGLEY FIELD, VA.\n\nCreated by act of Congress approved March 3, 1915, for the supervision and direction of the scientific study of the problems of flight (U. S. Code, Title 50, Sec. 151). Its membership was increased to 15 by act approved March 2, 1929. The members are appointed by the President, and serve as such without compensation.\n\nJOSEPH S. AMES, Ph. D., *Chairman*,\nBaltimore, Md.\n\nDAVID W. TAYLOR, D. Eng., *Vice Chairman*,\nWashington, D. C.\n\nCHARLES G. ABBOT, Sc. D.,\nSecretary, Smithsonian Institution.\n\nLYMAN J. BRIGGS, Ph. D.,\nDirector, National Bureau of Standards.\n\nARTHUR B. COOK, Rear Admiral, United States Navy,\nChief, Bureau of Aeronautics, Navy Department.\n\nWILLIS RAY GREGG, B. A.,\nChief, United States Weather Bureau.\n\nHARRY F. GUGGENHEIM, M. A.,\nPort Washington, Long Island, N. Y.\n\nSYDNEY M. KRAUS, Captain, United States Navy,\nBureau of Aeronautics, Navy Department.\n\nCHARLES A. LINDBERGH, LL. D.,\nNew York City.\n\nWILLIAM P. MACCRACKEN, JR., LL. D.,\nWashington, D. C.\n\nAUGUSTINE W. ROBINS, Brigadier General, United States Army,\nChief Matériel Division, Air Corps, Wright Field, Dayton, Ohio.\n\nEUGENE L. VIDAL, C. E.,\nDirector of Air Commerce, Department of Commerce.\n\nEDWARD P. WARNER, M. S.,\nNew York City.\n\nOSCAR WESTOVER, Major General, United States Army,\nChief of Air Corps, War Department.\n\nORVILLE WRIGHT, Sc. D.,\nDayton, Ohio.\n\nGEORGE W. LEWIS, *Director of Aeronautical Research*\n\nJOHN F. VICTORY, *Secretary*\n\nHENRY J. E. REID, *Engineer in Charge, Langley Memorial Aeronautical Laboratory, Langley Field, Va.*\n\nJOHN J. IDE, *Technical Assistant in Europe, Paris, France*\n\n## TECHNICAL COMMITTEES\n\nAERODYNAMICS\nPOWER PLANTS FOR AIRCRAFT\nAIRCRAFT STRUCTURES AND MATERIALS\n\nAIRCRAFT ACCIDENTS\nINVENTIONS AND DESIGNS\n\n*Coordination of Research Needs of Military and Civil Aviation*\n*Preparation of Research Programs*\n*Allocation of Problems*\n*Prevention of Duplication*\n*Consideration of Inventions*\n\n## LANGLEY MEMORIAL AERONAUTICAL LABORATORY\nLANGLEY FIELD, VA.\n\nUnified conduct, for all agencies, of scientific research on the fundamental problems of flight.\n\n## OFFICE OF AERONAUTICAL INTELLIGENCE\nWASHINGTON, D. C.\n\nCollection, classification, compilation, and dissemination of scientific and technical information on aeronautics.\n\nII\n```", "timestamp": "2026-07-19T18:22:05.858647+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 6, "total_pages": 16, "image_filename": "19930094559_p6.jpg", "text": "4 N.A.C.A. Technical Memorandum No. 857\n\ngive the total times from start of injection to end of combustion. Consideration of each of the processes leads to the conclusion that, in addition to the atomization of the fuel by the nozzle and further distribution and heating by the air, the fuel is further prepared for combustion by the ignition-lag intervals that become available. Good fuel preparation before combustion and the presence of sufficient combustion air quantities reduce the combustion time.\n\nThe shortest total time of ignition lag and combustion is attained when the air velocity during the fuel preparation interval is high. Since the energy of the air vortex in the chamber is greatest when the inflowing air velocity is greatest, the combustion in the chamber will then occur under the most favorable conditions. With start of injection taking place before maximum velocity of inflow, the time for complete combustion remains practically unchanged since the effect of the decreasing ignition lag — that is, shorter time for preparation of the fuel — and the increasing air velocity evidently offset each other. With later injection there is a strong increase in the time taken for combustion, since the interval of fuel preparation becomes still smaller and there is, moreover, a strong slowing down of the air movement. The greater density resulting from the higher pressure is not at all sufficient to offset this effect.\n\nFinally, for different fuel quantities, there were determined the combustion time-intervals, with the other conditions remaining the same (fig. 10). Within the practical range the combustion time increases approximately in proportion to the fuel quantity injected. In addition to the curve for the nozzle DN 12 SD 12, there are also given on the figure the results obtained under the same conditions, using the Bosch nozzle DN 30 S 2. For the latter nozzle the tests gave curves quite similar to those of figure 9, but the combustion time-intervals were longer, as may be seen on figure 10. This nozzle gives a spatially wider fuel distribution since the nozzle has a spray angle of $30^\\circ$, and a diameter of 2 mm as compared with the $12^\\circ$ angle and 0.5 mm diameter of the DN 12 SD 12 nozzle. The flame photographs which, due to space limitation, are not given here, showed both a change in spray form and a change in the combustion, which appeared less concentrated locally than that shown in the photographs here reproduced.", "timestamp": "2026-07-19T18:22:06.083481+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 9, "total_pages": 76, "image_filename": "19930094549_p9.jpg", "text": "N.A.C.A. Technical Memorandum No. 867\n\nwith\n\n$\\rho_{1} = 2 \\sqrt{A^{2} + B^{2}}$\n\n$\\sin \\varphi_{1} = \\frac{2A}{\\rho_{1}}$\n\n$\\cos \\varphi_{1} = \\frac{-2B}{\\rho_{1}}$\n\nSimilarly, the part of the root $l_{1} C_{1} e^{\\lambda_{1} t} + l_{2} C_{2} e^{\\lambda_{2} t}$ will assume the following forms:\n\n$e^{at} \\left[ (l_{1} C_{1} + l_{2} C_{2}) \\cos bt + i (l_{1} C_{1} - l_{2} C_{2}) \\sin bt \\right]$\n\n$e^{at} \\left[ (2A \\alpha - 2B \\beta) \\cos bt - (2A \\beta + 2B \\alpha) \\sin bt \\right]$\n\n$e^{at} \\rho_{2} \\sin (bt + \\varphi_{2})$\n\nwith\n\n$\\rho_{2} = 2 \\sqrt{(A\\alpha - B\\beta)^{2} + (A\\beta + B\\alpha)^{2}}$\n\n$\\sin \\varphi_{2} = \\frac{2(A\\alpha - B\\beta)}{\\rho_{2}}$\n\n$\\cos \\varphi_{2} = \\frac{-2(A\\beta + B\\alpha)}{\\rho_{2}}$\n\nWe may note that $\\rho_{2}$ may be written:\n\n$\\rho_{2} = 2 \\sqrt{(A^{2} + B^{2})(\\alpha^{2} + \\beta^{2})}$\n\nWe see that the values of the amplitudes $\\rho$ and of the phase displacements $\\varphi$ depend on the initial conditions. For the same oscillation the ratios $\\rho_{2}/\\rho_{1}$, $\\rho_{3}/\\rho_{1}$, $\\rho_{4}/\\rho_{1}$ are, however, independent of the initial disturbance and depend only on the characteristics of the airplane.", "timestamp": "2026-07-19T18:22:07.163633+00:00"}
{"citation_id": "19930094534", "source_url": "https://ntrs.nasa.gov/api/citations/19930094534/downloads/19930094534.pdf", "page_number": 14, "total_pages": 21, "image_filename": "19930094534_p14.jpg", "text": "N.A.C.A. Technical Memorandum No. 882 13\n\nV. BEAM STRAP MILLING MACHINE\n\nThe butt straps of beams usually wedge-shaped for statical reasons can, of course, be fabricated as a rolling mill product, but such ready rolled plates are practically out of the question in experimental construction; first, because the restricted number of such plates required would make an order for rolling uneconomical, then too, delivery could not be had immediately. For this reason, the Heinkel Company had the firm of Paul Knopp, Berlin S W 19, develop and construct a milling machine which meets the requirements satisfactorily and makes the tedious and at the same time expensive hand work altogether superfluous. Plates up to 15 mm thickness and 150 mm width and of any length can be utilized on this machine, and any desired pitch from 1:∞ to about 1:20 can be milled with the mechanical speed regulator. For this purpose the feed screw is driven over a revolution regulator with adjustable speed. With this regulator the feed screw can run forward or backward or be locked. In the latter case, the milled plates will be of constant thickness. Direction of rotation with the cut gives taper, against the cut gives the opposite effect.\n\nThe milling cutter operates at 250 and at 450 m/min. cutting speeds. It has a two-speed feed. The higher stage 1 m/min. is for roughing down, the other for finishing. The thickness of the chip properly ranges between 0.5 and 0.8 mm for roughing. The feed rollers are interchangeable and can be exchanged for section rollers. For example, angle sections with shanks up to 30 mm can be milled.\n\nSamples from this machine are:\n\nDuraluminum fitting plate, 4 x 100 x 2,720, thinned down to 1.5 mm at both ends for a distance of 200 mm;\n\nA 15 x 126 x 4,112 mm plate of the same metal thinned down over its entire length of 15 mm to 2,5 mm.\n\nThe finished product has an unobjectionably smooth surface and requires no finishing by hand.", "timestamp": "2026-07-19T18:22:12.665131+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 13, "total_pages": 47, "image_filename": "19930093641_p13.jpg", "text": "mains about the same for $\\beta = 28-1/2^\\circ$. The efficiency at the climb condition increases progressively with increasing blade angle up to $\\beta = 28-1/2^\\circ$. The effects of blade-angle setting for the enclosed-engine arrangement with tractor propellers at positions 2 and 3 are shown in figures 29 and 30. These data indicate in general that, up to a blade angle of about $28-1/2^\\circ$, the propulsive efficiencies remain substantially the same.\n\nA continuation of this investigation to cover a wider range of blade angles may be of interest, particularly for high-speed airplanes for which values of $\\beta = 40^\\circ$ are not uncommon. Values of the maximum propulsive efficiency for all arrangements are given in table I.\n\nValues of over-all efficiency computed by means of the previously defined formula are given in table I for all arrangements at lift coefficients of 0.25 and 0.70. For the lift coefficient corresponding to high speed, the pusher and tractor positions 1 and 2 have over-all efficiencies of 79 and 78 percent, respectively, whereas the model with the conventional wing nacelles has an over-all efficiency of 60 percent with exposed radiators and 72 percent without radiators. For the lift coefficient corresponding to the climb condition, the efficiencies vary from 78 percent for the best enclosed-engine arrangement to 68 percent for the wing-nacelle arrangement with radiators. No allowance has been made for radiator drag in the over-all efficiencies of the enclosed-engine arrangement.\n\nFaired spinners on the extension shafts appear to have a negligible effect on over-all efficiency. The overall efficiencies for tractor position 3 were definitely inferior, being 3 percent below those for tractor position 1 at the high-speed condition and 6 percent below at climb.\n\nPOWER-ON CHARACTERISTICS\n\nThe effect of power on the lift and pitching moments of the model for some of the test conditions is shown in figures 31 to 36. In the presentation of the results, the power-on condition for each test is denoted by the index thrust coefficient $T_{co}'$. This coefficient is defined by\n\n$$\nT_{co}' = \\frac{T - \\Delta D}{qS}\n$$", "timestamp": "2026-07-19T18:22:15.867576+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 24, "total_pages": 32, "image_filename": "19930094552_p24.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFigs. 3,5,6,12\n\n[Figure: Test set-up for bending loading condition]\nFigure 3.-Test set-up for bending\nloading condition\n\n[Figure: Application of forces on test specimen under the arching loading condition]\nFigure 6.- Application of forces\non test specimen under\nthe arching loading condition.\n\n[Figure: Test set-up for arching loading condition]\nFigure 5.- Test set-up for arching\nloading condition.\n\n[Figure: Bulging of the cylinder under an arching load]\nFigure 12.- Bulging of the cylinder\nunder an arching load.", "timestamp": "2026-07-19T18:22:29.841544+00:00"}
{"citation_id": "19930091716", "source_url": "https://ntrs.nasa.gov/api/citations/19930091716/downloads/19930091716.pdf", "page_number": 17, "total_pages": 24, "image_filename": "19930091716_p17.jpg", "text": "```markdown\nNEGATIVE THRUST AND TORQUE OF SEVERAL FULL-SCALE PROPELLERS 13\n\nand there is evidence to show that this deviation becomes greater as the temperature decreases. Fortunately, however, the deviation from the linear formula in the range of low engine speeds usually occurs in a range of blade angles where large changes in torque cause but small changes in thrust coefficient and also where the values of thrust coefficients are low so that, although the relative error may be large, the absolute value of the error is small.\n\nIt is also seen that, at high values of engine speed (above 2,000 r. p. m.), the friction torque increases faster than the straight-line assumption, so that, in problems involving high values of $N_e$ (fast dives), it will be advisable to increase $Q_f/N_e$ by a small amount.\n\n**Estimation of a value of $Q_f/N_e$.**—On the assumption that the friction torque can be determined by the equation $Q_f=\\Delta N_e K$, it is clear that a plot of $Q_f/N_e$ against displacement $\\Delta$ for a group of engines should be a straight line like the dotted line, taken from figure 5 of reference 3, in the lower plot in figure 14.\n\nStraight lines were drawn through the friction torque curves, as illustrated for engine 3 (fig. 14), and the values of $Q_f/N_e$ represented by these lines were plotted against $\\Delta$ in the lower chart. The solid faired line through these points may possibly provide a more accurate selection of $Q_f/N_e$ values than the broken line from reference 3.\n\nWhere no specific friction data are available, a reasonable estimate of the value of $Q_f/N_e$ for any engine, may be obtained from this curve. In the case of a geared engine, $Q_f/N_e$ must be converted to $Q_f/N$ when calculating the coefficient $Q_c$.\n\n**Applicability of friction data.**—The friction torque of engines varies with many factors; such as the mechanical condition of the engine, the cylinder barrel and oil temperatures, the barometric pressure, the throttle opening, the oil viscosity, and the gear ratio.\n\nIn the selection of a value of $Q_f/N_e$ from figure 14, some allowance should properly be made for these factors. Some of the factors, however, tend to cancel each other; some, such as mechanical condition, have an unpredictable effect; and others have but a small effect. In general, there will probably be little justification for making any corrections but, under extreme conditions, these factors should not be overlooked.\n\nThe effect of altitude is to reduce the friction torque (pumping losses), but this gain is balanced by the increased friction due to the lower temperatures existing at the higher altitudes. Corrections for altitude are therefore unnecessary in most cases. Gearing an engine should not alter the friction torque by more than 10 to 20 percent at rated engine speed. Changes in temperature will have a considerable effect and may change the specific friction torque $Q_f/\\Delta$ by as much as 0.004 per 10° F. change in outside-air temperature. It is beyond the scope of this report to consider in detail the quantitative aspects of the effects of the many factors that affect the friction torque of engines. Considerable information of this nature is given in reference 6.\n\n**APPLICATION OF NEGATIVE THRUST AND TORQUE DATA**\n\nThe development of the controllable propeller has greatly increased the opportunities for using the negative thrust of a propeller to advantage or, in other instances, for avoiding the bad effects of negative thrust at one blade angle by changing to another angle where these effects are less severe. Some of the ways in which negative thrust and torque data may be used to deal with such problems are given in the following paragraphs.\n\n**DRAG OF PROPELLER ON DEAD ENGINE OF A MULTIENGINE AIRPLANE**\n\nOne problem of interest to the operators of multiengine airplanes concerns the question of flying with one or more engines dead. In this situation it is necessary to reduce the drag of the airplane to a minimum so that the power of the remaining engines will be sufficient to maintain the altitude required to clear all obstacles on the path to the nearest airport. It is of considerable interest, therefore, to know just where to set the blade angle of the dead-engine propeller to absorb the least power. Such problems may be readily solved by the data given in this report.\n\n**Example.**—An example of one such problem is carried through to show the method of attack. The assumed conditions are as follows:\n\nAirplane flying at 135 miles per hour with one engine dead.\n\nEngines (2)—750 horsepower; $\\Delta=1,500$ cubic inches; $N=1,450$; $N_e=2,000$.\n\nPropellers—R. A. F. 6 section; 3 blades; 11-foot diameter.\n\nAltitude—5,000 feet; $\\sigma=0.862$; $Q_f/N_e=-0.09$ from figure 14; $Q_f/N=-0.1715$ and, after adding 10 percent for gearing, becomes -0.1885.\n\n$$Q_c=17,200 \\times \\frac{Q_f/N}{\\sigma D^5 V^2} = \\frac{17,200 \\times -0.1885}{0.862 \\times 11^5 \\times 135^2} = -0.0019.$$\n\nIn figure 7 the line representing $Q_c$ is drawn in (line O-C) and its intersection with the $Q_c$ curve for any blade angle represents the value of $Q_c$ for that particular blade angle and the corresponding value of $T_c$ may be obtained by projecting up from this intersection to the $T_c$ curve for the corresponding blade angle (line D-D').\n\nAs previously pointed out, the assumption that the friction torque approaches zero at low values of $N_e$ (low values of $nD/V$) does not hold very well, though the absolute value of the error resulting from this assumption is small. As an added refinement, this error may be reduced as follows:\n\nEstimate the value of friction torque at $N_e=0$ and from it calculate a value of $Q_c$ at which the propeller will stop turning. From that point on the $Q_c$ ordinate\n```", "timestamp": "2026-07-19T18:22:34.670275+00:00"}
{"citation_id": "19930094543", "source_url": "https://ntrs.nasa.gov/api/citations/19930094543/downloads/19930094543.pdf", "page_number": 43, "total_pages": 50, "image_filename": "19930094543_p43.jpg", "text": "N.A.C.A. Technical Memorandum No. 873\n\nFigure 6.- Vertical section of cylinder head\nof form No. 4.\n\nFigure 8.- Vertical section of head of form\nNo. 3.\n\nFigure 7.- Form of head Nos. 3,4 or 5 seen\nfrom below.\n\nFigure 9.- Vertical section of head of form\nNo. 5.\n\nFigs. 6,7,8,9", "timestamp": "2026-07-19T18:22:42.592667+00:00"}
{"citation_id": "19930094564", "source_url": "https://ntrs.nasa.gov/api/citations/19930094564/downloads/19930094564.pdf", "page_number": 6, "total_pages": 16, "image_filename": "19930094564_p6.jpg", "text": "N.A.C.A. Technical Memorandum No. 852 5\n\nment chamber which for low turbulence corresponds to the DVL tunnel and, as regards the obtainable Reynolds Numbers, extends the DVL measurements in the direction of lower Reynolds Numbers. The comparison with this tunnel was to prove the reliability of the DVL tunnel at low speeds. Unfortunately, only one test series of this tunnel is known (reference 7).\n\nThe N.A.C.A. VDT is a 1.5-meter high-pressure tunnel with closed test section whose jet is very turbulent (T.F. = 2.64). This tunnel was included in our comparison because with its effective Reynolds Numbers it extends the DVL measurements toward larger R and at the time is the tunnel in which the most extensive systematic measurements have been made so far. With its maximum attainable Reynolds Number the N.A.C.A. VDT offers any amount of desirable data. On the other hand, only very little data on systematic tests with low Reynolds Numbers are available (references 3 and 8). The N.A.C.A. VDT data at low Reynolds Number have not been included in the figures 3 and 8 because they are not systematic and apparently disclose scattering. Adding this scarce material would prove nothing while detracting from the otherwise lucid representation.\n\nAnalysis of the entire data in figures 3 to 5 manifest the following:\n\n1) The DVL findings agree with the unfortunately scarce result of the GALCIT (on airfoil section 2412). For effective Reynolds Number $\\cong 1.5 \\times 10^5$ the result in the DVL tunnel seems to be more reliable than that in the GALCIT. The somewhat too small results of the latter at its maximum Reynolds Numbers are probably due to the fact that at maximum speed the wire-suspended models are readily somewhat disturbed and consequently give slightly lower maximum lift. This effect is probably also the cause for various identical deviations in the DVL measurements (airfoil sections 2418 and 2421).\n\n2) With exception of sections 2418 and 2421, the extrapolation of the DVL measurements joins the results of the N.A.C.A. VDT satisfactorily so far as the VDT results for maximum pressure (effective $R \\cong 8 \\times 10^6$) are used. There", "timestamp": "2026-07-19T18:22:42.768275+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 14, "total_pages": 47, "image_filename": "19930093641_p14.jpg", "text": "```markdown\n12\n\nand is nondimensional and similar to a drag coefficient. In order to determine the $T_{c_o}'$ corresponding to a given operating condition of the propeller, it was found convenient to replace the effective thrust $T - \\Delta D$ by its equivalent $P\\eta/V$, where $P$ is the total power to all the propellers. Since $\\eta$ varies only slightly with lift coefficient, it was arbitrarily replaced by $\\eta_o$, the propulsive efficiency at $C_L = 0.25$, so that\n\n$$T_{c_o}' = \\frac{P\\eta_o}{qSV}$$\n\nThe variations of lift coefficient with $T_{c_o}'$ for the pusher and tractor position 1 are shown in figures 31 and 32. In both cases the tests were made at a tunnel speed of approximately 30 miles per hour in order to reach large values of $T_{c_o}'$ with the available power. The effect of power in both conditions is similar in that the lift-curve slope and the maximum lift coefficient are increased in almost a linear fashion with increasing values of $T_{c_o}'$ (fig. 33). The effect of power is more pronounced for the tractor-propeller condition, inasmuch as the slipstream velocity over the wing is higher than the inflow velocity for the pusher propellers. Computations indicate that part of the increased lift from the pusher propellers is obtained from boundary-layer control by delaying separation at the trailing edge of the wing.\n\nIn figures 34, 35, and 36 the pitching-moment coefficients for the model with conventional wing nacelles, the model pusher, and tractor position 1 are shown over a range of values of $T_{c_o}'$. The pusher is superior to both of the tractor arrangements with respect not only to greater static stability at the high-speed conditions but also to smaller changes in balance with increasing power. Power has a generally similar effect on the pitching-moment coefficients of tractor position 1 and the wing-nacelle arrangement.\n\nPROPELLER NOISE\n\nInasmuch as the choice of propeller positions will to some extent be governed by the propeller noise, the meas-\n```", "timestamp": "2026-07-19T18:23:11.623556+00:00"}
{"citation_id": "19930094549", "source_url": "https://ntrs.nasa.gov/api/citations/19930094549/downloads/19930094549.pdf", "page_number": 10, "total_pages": 76, "image_filename": "19930094549_p10.jpg", "text": "8\nN.A.C.A. Technical Memorandum No. 867\n\nAIRPLANES STUDIED\n\nWe shall study a series of airplanes differing only\nin the coefficient of static stability $\\partial C_M/\\partial i$. If $i$\nmay be taken equal to $-\\frac{w}{V}$, the derivatives with respect\nto $w$ figuring in the calculations are merely the deriva-\ntives with respect to the angle of attack $i$ multiplied\nby $-\\frac{1}{V}$. We have shown elsewhere that the longitudinal\nmotion of an airplane is sufficiently determined by: five\ngeometric characteristics; seven aerodynamic characteris-\ntics; a characteristic depending on the engine propeller\nunit; and a characteristic depending on the weight. In\nour example, we have assumed the following values for\nthese elements:\n\n1. Geometric characteristics:\n   $l = 2.3$ meters\n   $r = 1.3$ \"\n   $s = 0$\n   $l' = 2.6 \\ l$\n   $S' = \\frac{1}{7} \\ S$\n\n$l'$ and $S'$ being the lever arm and the area of the horizon-\ntail surface, respectively.\n\n2. Aerodynamic characteristics:\nThe quantities $C_x$ and $C_z$ are the coefficients of\nthe aerodynamic forces in the directions of the\naxes fixed to the wing.\n\nThe angle of attack $i$ of the state considered is 0:\n$C_z = 0.40$.\n\nMinimum drag-lift ratio at this angle of attack:\n$\\beta = 0.125$.\n\n$\\frac{dC_x}{di} = 0.006$; $\\frac{dC_z}{di} = 0.065$.", "timestamp": "2026-07-19T18:23:17.839567+00:00"}
{"citation_id": "19930091692", "source_url": "https://ntrs.nasa.gov/api/citations/19930091692/downloads/19930091692.pdf", "page_number": 12, "total_pages": 20, "image_filename": "19930091692_p12.jpg", "text": "8\nREPORT NO. 617—NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS\n\n[Figure: Three oscillograph traces labeled 168, 161, and 156. Vertical axis: PRESSURE, LB. PER SQ. IN. Horizontal axis: TIME, SEC.]\n\n| Record | Air-fuel ratio | Density (lb./cu. ft.) | Ignition lag (sec.) | Maximum explosion pressure (lb./sq. in.) |\n| :--- | :--- | :--- | :--- | :--- |\n| 168 | 20 | 0.59 (air) | 0.0018 | 800 |\n| 161 | 20 | 0.59 (air) | .0020 | 960 |\n| | | 0.29 (nitrogen) | | |\n| 156 F | 45 | 0.89 (air) | .0013 | 860 |\n| 156 | 20 | 0.59 (air) | .0018 | 970 |\n| | | 0.32 (products) | | |\n\n[Figure: Three oscillograph traces labeled 154, 147, and 142. Vertical axis: PRESSURE, LB. PER SQ. IN. Horizontal axis: TIME, SEC.]\n\n| Record | Air-fuel ratio | Density (lb./cu. ft.) | Ignition lag (sec.) | Maximum explosion pressure (lb./sq. in.) |\n| :--- | :--- | :--- | :--- | :--- |\n| 154 | 20 | 0.89 (air) | 0.0013 | 1,260 |\n| 147 | 20 | 0.89 (air) | .0014 | 1,480 |\n| | | 0.28 (nitrogen) | | |\n| 142 F | 60 | 1.18 (air) | .0012 | 1,150 |\n| 142 | 20 | 0.89 (air) | .0018 | 1,420 |\n| | | 0.31 (products) | | |\n\nFIGURE 10.—Effect of inert gases on ignition and combustion with a gas temperature of 1,155° F.", "timestamp": "2026-07-19T18:23:19.108836+00:00"}
{"citation_id": "19930094552", "source_url": "https://ntrs.nasa.gov/api/citations/19930094552/downloads/19930094552.pdf", "page_number": 25, "total_pages": 32, "image_filename": "19930094552_p25.jpg", "text": "N.A.C.A. Technical Memorandum No. 864\nFigs. 4,7,8\n\nSection C-D/Section E-F\nSection A-B\n[Figure: Diagram of Angle, Lining piece, and Section]\nFigure 4.- Application of forces on main spars.\n\nSheet\n[Graph: Prebuckling range load 1738 kg]\n(a) Prebuckling range load 1738 kg\n[Graph: Buckling range load 4475 kg]\n(b) Buckling range load 4475 kg\n\nSketch of measuring sections\n[Graph: Stress distribution over the cross section height under the bending loading.]\nFigure 7.- Stress distribution over the cross section height under the bending loading.\n\n[Graph: Variation of the stringer forces over the cylinder length for the bending loading.]\nFigure 8.- Variation of the stringer forces over the cylinder length for the bending loading.", "timestamp": "2026-07-19T18:23:24.165399+00:00"}
{"citation_id": "19930094559", "source_url": "https://ntrs.nasa.gov/api/citations/19930094559/downloads/19930094559.pdf", "page_number": 7, "total_pages": 16, "image_filename": "19930094559_p7.jpg", "text": "N.A.C.A. Technical Memorandum No. 857 5\n\nInsofar as a rapid combustion process is considered to be important, a distribution of the fuel over a larger air space than that corresponding to the local combustion air requirement does not appear to be justified. The speed of combustion will in that case become smaller since the effect of each partial combustion on the successive combustion of the neighboring particles is reduced as a result of the greater spatial separation.\n\n3. TESTS ON MODEL ENGINE OF THE AIR-STORAGE OR AIR-CELL TYPE\n\nIn order to observe the movement of the fuel during the ignition lag, the scavenging of the air-storage chamber as well as the combustion process itself, the model shown in figures 11 and 12 was built into the combustion bomb, the design of the model following that of the Henschel-Lanova engine type K. The distance of the nozzle from the air chamber, the arrangement of the air spaces, and approximate sizes and cross sections of the throttles were made to correspond to the engine. The fuel is sprayed through the main combustion space and into two air spaces or coils lying one behind the other and communicating with each other and with the combustion space through throttle passages. The two upper windows in the model provide means for observation into the main combustion chamber. The first air coil is provided with a small window which enables the start of ignition at this position to be followed. The second air cell is provided with two windows, making it possible to observe the quantity of unignited fuel that has reached this position and how the combustion is there developed. The air flows through six orifices on one side of the bomb into the main combustion space, following the same direction as in the comparison engine. The Bosch nozzle DN 4 S9 and the opening pressure of 150 atm., correspond to the engine.\n\nTest results.- For the flame pictures shown in figures 13 and 14, the start of injection occurs approximately within the range of maximum velocity of inflow of air through the first throttle. The fuel quantities correspond to about one-quarter load for figure 13, and full load for figure 14. In the pictures obtained without air movement it was observed that the fuel was partially kept back without motion in front of the mouth of the first", "timestamp": "2026-07-19T18:23:29.485351+00:00"}
{"citation_id": "19930091715", "source_url": "https://ntrs.nasa.gov/api/citations/19930091715/downloads/19930091715.pdf", "page_number": 29, "total_pages": 30, "image_filename": "19930091715_p29.jpg", "text": "Positive directions of axes and angles (forces and moments) are shown by arrows\n\n| Axis | | Force (parallel to axis) symbol | Moment about axis | | | Angle | | Velocities | |\n| :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- | :--- |\n| Designation | Symbol | | Designation | Symbol | Positive direction | Designation | Symbol | Linear (component along axis) | Angular |\n| Longitudinal<br>Lateral<br>Normal | $X$<br>$Y$<br>$Z$ | $X$<br>$Y$<br>$Z$ | Rolling<br>Pitching<br>Yawing | $L$<br>$M$<br>$N$ | $Y \\longrightarrow Z$<br>$Z \\longrightarrow X$<br>$X \\longrightarrow Y$ | Roll<br>Pitch<br>Yaw | $\\phi$<br>$\\theta$<br>$\\psi$ | $u$<br>$v$<br>$w$ | $p$<br>$q$<br>$r$ |\n\nAbsolute coefficients of moment\n$C_l = \\frac{L}{qbS}$ (rolling)\n$C_m = \\frac{M}{qcS}$ (pitching)\n$C_n = \\frac{N}{qbS}$ (yawing)\n\nAngle of set of control surface (relative to neutral position), $\\delta$. (Indicate surface by proper subscript.)\n\n4. PROPELLER SYMBOLS\n\n$D$, Diameter\n$p$, Geometric pitch\n$p/D$, Pitch ratio\n$V$, Inflow velocity\n$V_s$, Slipstream velocity\n$T$, Thrust, absolute coefficient $C_T = \\frac{T}{\\rho n^2 D^4}$\n$Q$, Torque, absolute coefficient $C_Q = \\frac{Q}{\\rho n^2 D^5}$\n\n$P$, Power, absolute coefficient $C_P = \\frac{P}{\\rho n^3 D^5}$\n$C_s$, Speed-power coefficient $= \\sqrt[5]{\\frac{\\rho V^5}{P n^2}}$\n$\\eta$, Efficiency\n$n$, Revolutions per second, r.p.s.\n$\\Phi$, Effective helix angle $= \\tan^{-1} \\left( \\frac{V}{2\\pi r n} \\right)$\n\n5. NUMERICAL RELATIONS\n\n1 hp. = 76.04 kg-m/s = 550 ft-lb./sec.\n1 metric horsepower = 1.0132 hp.\n1 m.p.h. = 0.4470 m.p.s.\n1 m.p.s. = 2.2369 m.p.h.\n\n1 lb. = 0.4536 kg.\n1 kg = 2.2046 lb.\n1 mi. = 1,609.35 m = 5,280 ft.\n1 m = 3.2808 ft.\n\n27", "timestamp": "2026-07-19T18:23:32.735854+00:00"}
{"citation_id": "19930091707", "source_url": "https://ntrs.nasa.gov/api/citations/19930091707/downloads/19930091707.pdf", "page_number": 17, "total_pages": 20, "image_filename": "19930091707_p17.jpg", "text": "CRINKLING STRENGTH AND BENDING STRENGTH OF ROUND AIRCRAFT TUBING 13\n\nThe modulus of rupture, in nondimensional form, of the chromium-molybdenum-steel tubes was found to be given by\n\n$$\n\\sigma_{rs} = \\frac{1.525 \\frac{1}{\\delta_y}}{\\frac{1}{\\delta_y} + 1.4}, \\quad \\frac{1}{\\delta_y} > 5\n$$\n\n(8)\n\nThe modulus of rupture, in nondimensional form, of the duralumin tubes was found to be given by\n\n$$\n\\sigma_{rd} = 1.773 - \\frac{1.7}{\\frac{1}{\\delta_y}}, \\quad \\frac{1}{\\delta_y} > 2.5\n$$\n\n(9)\n\nFigures 9 and 11 show also the theoretical curve (straight line) for axially symmetrical elastic failure.\n\nIf one assumes, for convenience, $\\mu = \\frac{1}{3} \\sqrt{\\frac{2}{3}} = 0.2722$, one obtains from equation (3)\n\n$$\n\\sigma_{cr} = \\frac{6 \\sqrt{\\tau}}{5 \\delta}\n$$\n\n(10)\n\nand, since for elastic failure, $\\tau = 1$\n\n$$\n\\sigma_{cr} = \\frac{6}{5 \\delta}\n$$\n\n(11)\n\nWhen $\\sigma_{cr}$ is plotted against $1/\\delta$, equation (11) may be represented by the straight line shown in the figures. It immediately appears that, by substituting for $\\tau$ in equation (10) the expression found from column tests, or otherwise, a theoretical relation between $\\delta$ and $\\sigma_{cr}$ might be obtained in the plastic range. This substitution might be done for the range of values of $\\sigma$ for which an expression for $\\tau$ was reliable; but no agreement with the results of crinkling tests would be expected because crinkling failures obtained in the laboratory in the plastic range are not stability failures but bending failures. This condition is necessarily true because of the impossibility of satisfying the end and other conditions required by theory for a stability failure.\n\nIt remains to obtain expressions for the crinkling strength and the bending strength of tubing that just complies with specifications. Such expressions may be obtained immediately from equations (2) and equations (4) to (9), inclusive. The specified minimum yield strength of chromium-molybdenum-steel tubing such as used in this investigation is, according to Navy Department Specification 44T18c for tubing not over 0.188 inch thick, 75,000 pounds per square inch, and the modulus of elasticity may be taken as 29,800,000 pounds per square inch (reference 4). Substituting these values in equations (5) and then replacing $\\delta_y$ and $\\sigma_{cry}$ in equation (6) by the expressions obtained from equations (5), and solving for the crinkling strength, $f_{cr}$, gives after rounding off,\n\n$$\nf_{cr} = \\frac{19\\,800\\,000}{\\frac{d}{t} + 200}, \\quad \\frac{d}{t} < 80\n$$\n\n(12)\n\nin lb per sq in.\n\nThe specified minimum tensile yield strength of duralumin tubing such as used in this investigation is, according to Navy Department Specification 44T21b for Condition “T” heat-treated tubing, 40,000 pounds per square inch. The average ratio of compressive yield strength to tensile yield strength of the tubes used in this investigation was found to be 0.864. The average value of the modulus of elasticity was 10,610,000 pounds per square inch. Substituting $S = 0.864 \\times 40,000 = 34,560$ pounds per square inch and $E = 10,610,000$ pounds per square inch in equations (2) and then replacing $\\delta_y$ and $\\sigma_{cry}$ in equation (7) by the expressions obtained from equations (2), and solving for the crinkling strength, $f_{cr}$, gives, after rounding off,\n\n$$\nf_{cr} = 56\\,400 \\left[ 1 - \\frac{\\frac{d}{t} - 1}{150 + \\frac{4}{3} \\left( \\frac{d}{t} - 1 \\right)} \\right], \\quad \\frac{d}{t} < 125\n$$\n\n(13)\n\nin lb per sq in.\n\nAn expression for the modulus of rupture of chromium-molybdenum-steel tubing that just complies with Navy Department Specification 44T18c for tubing not over 0.188 inch thick may be found, as just outlined, from equations (5) and (8):\n\n$$\nf_r = \\frac{32\\,500\\,000}{\\frac{d}{t} + 283}, \\quad \\frac{d}{t} > 80\n$$\n\n(14)\n\nin lb per sq in.\n\nAn expression for the modulus of rupture of duralumin tubing that just complies with Navy Department Specification 44T21b for Condition “T” heat-treated tubing may be found, as outlined, from equations (4) and (9):\n\n$$\nf_r = 61\\,500 - 191 \\frac{d}{t}, \\quad \\frac{d}{t} < 125\n$$\n\n(15)\n\nin lb per sq in.\n\nThe curves representing equations (12), (13), (14), and (15) are shown in figures 8, 10, 12, and 14, respectively. They indicate, for the range of values of ratio of diameter to thickness covered, the crinkling strengths and the moduli of rupture that may be expected from tubing which just complies with the applicable specifications noted.\n\nDISCUSSION\n\nAs explained in the introduction, the crinkling strength is the upper limit of column strength. With the determination of the crinkling strength, it now becomes possible to indicate where the column curves for the two materials of this investigation must be “cut off” at their upper ends for tubing of a given ratio of diameter to thickness, namely, at the stresses given by equations (12) and (13).\n\n* The two points in figure 12 that are below the curve represent the results of tests on specimens the material of which did not comply with the requirement of the specification for chemical composition per yield strength (ICS-T, fig. 3).", "timestamp": "2026-07-19T18:23:34.013519+00:00"}
{"citation_id": "19930091655", "source_url": "https://ntrs.nasa.gov/api/citations/19930091655/downloads/19930091655.pdf", "page_number": 5, "total_pages": 22, "image_filename": "19930091655_p5.jpg", "text": "REPORT No. 580\n\nHEAT TRANSFER TO FUEL SPRAYS INJECTED INTO HEATED GASES\n\nBy ROBERT F. SELDEN and ROBERT C. SPENCER\n\nSUMMARY\n\nA study has been made of the influence of several variables on the pressure decrease accompanying injection of a relatively cool liquid into a heated compressed gas. Indirectly, this pressure decrease and the time rate of change of it are indicative of the total heat transferred as well as of the rate of heat transfer between the gas and the injected liquid. Air, nitrogen, and carbon dioxide were used as ambient gases; Diesel fuel and benzene were the injected liquids. The gas densities and gas-fuel ratios covered approximately the range used in compression-ignition engines. The gas temperatures ranged from $150^\\circ$ C. to $350^\\circ$ C. Several general conclusions may be drawn from the experimental results: Vaporization begins immediately after the start of injection; the initial rate of heat transfer is a direct function of the initial temperature difference between the gas and the fuel; and the heat transfer is less efficient the greater the injected fuel quantity, even though the total heat transferred is greater.\n\nINTRODUCTION\n\nIt is generally recognized that the compression-ignition engine in its present state of development suffers the disadvantage of inefficient utilization of its air charge. Recognizing that the utilization of the air must be partly dependent upon the fuel spray, Lee has conducted a detailed photographic investigation of the exterior characteristics of fuel sprays (reference 1). He has also determined the spatial distribution of the fuel within the spray (reference 2). These spray investigations have been extended by tests with the N. A. C. A. combustion apparatus and the results give an improved insight into the gross physical and chemical processes as they occur in the engine. (See references 3 and 4.)\n\nThe ignition lag in compression-ignition engines has been shown to influence the character of the subsequent explosion (references 5 and 6). No entirely satisfactory explanation of this fact has been given, but certain general conclusions can be drawn: In general, the fuel must be heated after injection; the fuel and the air must be mixed; and certain preliminary chemical reactions must take place before the actual ignition can occur. The observed lag is thus a composite of the intervals associated with these processes. It follows that heating the fuel prior to injection cannot reduce the ignition lag indefinitely although some reduction may be accomplished in this manner (reference 7). Rothrock and Waldron have shown that appreciable vaporization follows injection of the fuel into the combustion apparatus (reference 8). The time required for this vaporization to begin was not established but, in view of Wentzel’s theoretical analysis of the heating and vaporization of fuel droplets suspended in a heated gas (reference 9), there is every reason to believe that appreciable vaporization occurs in a compression-ignition engine during the ignition-lag period.\n\nThe present investigation was undertaken to isolate the heat transfer accompanying the mixing of a fuel spray and the ambient gas in a bomb and to study the influence of several variables on this individual process. The results of this investigation should give an insight into the time required to effect some vaporization since this process necessarily corresponds to a portion of the total heat transfer. Experimentally, heat transfer is not directly measurable in a system of this type; therefore, resort has been had to an indirect approach, namely, the measurement of the change in pressure accompanying the adiabatic exchange of heat between the gas and fuel after injection of the latter into a bomb. The primary variables were the gas temperature, the gas density, and the gas-fuel ratio. The effects of the nozzle design, the fuel temperature, the kind of fuel, and the character of the ambient gas were less extensively investigated.\n\nGas densities covering most of the range found in engine practice were used. For mechanical reasons temperatures corresponding to those attained in compression-ignition engines at top center could not be used. The maximum temperature employed was actually somewhat less than that of the gas charge prevailing at the start of injection (references 10 and 11) in compression-ignition engines.\n\nANALYSIS OF THE PROBLEM\n\nThe transfer of heat to a suspended droplet can take place by two mechanisms: conduction and radiation. Except insofar as their boundary conditions are altered, the mass flow of gas, induced by the injection of the liquid fuel, presumably is of little importance with respect to the individual droplets because of their low relative velocity (reference 12). The situation may be", "timestamp": "2026-07-19T18:23:41.729209+00:00"}
{"citation_id": "19930094538", "source_url": "https://ntrs.nasa.gov/api/citations/19930094538/downloads/19930094538.pdf", "page_number": 31, "total_pages": 43, "image_filename": "19930094538_p31.jpg", "text": "N.A.C.A. Technical Memorandum No. 878 29\n\nmental angles of twist would give the same results. With these free values, the breaking torque is computed 8.1 percent too high; this discrepancy is of the order of magnitude of the scatter in the buckling-bending tests. On cylinder No. III, the values $R_0$ and $\\delta_g$ were so chosen that the discrepancy remained small for the angle of twist as well as for the breaking torque. With a value $R_0 = 0.74$, the theoretical and experimental values of the angle of twist are identical, while the theoretical breaking torque is about 25 percent too high.\n\nIt should be remembered that the buckling-bending test with sheet panels does not completely reproduce the effect of the wrinkling deflections at the stringers, because the wrinkles — in contrast to the transverse load applied in panel tests — twist the section flange and so initiate its buckling.\n\nTranslation by J. Vanier,  \nNational Advisory Committee  \nfor Aeronautics.\n\n---\n\n**REFERENCES**\n\n1. Heck, O. S., and Ebner, Hans: Methods and Formulas for Calculating the Strength of Plate and Shell Constructions as Used in Airplane Design. T.M. No. 785, N.A.C.A., 1936.\n\n2. Wagner, H., and Ballerstedt, W.: Tension Fields in Originally Curved, Thin Sheets During Shearing Stresses. T.M. No. 774, N.A.C.A., 1935.", "timestamp": "2026-07-19T18:23:52.872690+00:00"}
{"citation_id": "19930094551", "source_url": "https://ntrs.nasa.gov/api/citations/19930094551/downloads/19930094551.pdf", "page_number": 8, "total_pages": 18, "image_filename": "19930094551_p8.jpg", "text": "N.A.C.A. Technical Memorandum No. 865\n\nIV. DESCRIPTION OF NOMOGRAPHIC METHOD EMPLOYED  \nTO DEFINE THE SUSPENSION-HOSE CURVE\n\nThere are several known methods of defining the form  \nof the curve resulting from towing heavy bodies from a  \nflexible cable through air for the purpose of predicting  \nthe position of the air-speed head in space. However,  \nthose methods are in part time-consuming, or, as is the  \ncase of Glauert’s method (reference 2), they proceed from  \nassumptions regarding the lift and drag of the hose, which  \nare considerably at variance with corresponding wind-tunnel  \ntests. For this reason, it seemed expedient to evolve a  \nmethod which conformed to actual conditions over a wide  \nfield, as well as being short. And the results as regards  \ndesirable accuracy fulfilled our expectations.\n\nThe method is predicated on a chord curve rather than  \nthe sag curve and postulates that the cable stiffness of  \nthe suspension has only a minor influence on the form of  \nthe curve itself owing to the comparatively great curvature radii of the curve.\n\nThe suspension hose is first divided into chord pieces  \nof equal finite length. Chord lengths of 50 cm proved  \nsmall enough to be serviceable as substitute for the related arc piece as theoretical basis.\n\nNotation\n\n$c_a$, lift coefficient of a hose element  \n$c_w$, drag coefficient of a hose element  \n$q$, flight dynamic pressure ($kg/m^2$)  \n$A_{1,2}$ vertical force component at hose element (kg)  \n$B_{1,2}$ horizontal force component at hose element (kg)  \n$\\Delta l$ length of suspension hose element (m)  \n$d$ diameter of suspension hose (m)  \n$\\Delta F = \\Delta l d$ area of suspension hose element ($m^2$)  \n$G_s$, weight of suspension hose element (kg)", "timestamp": "2026-07-19T18:23:57.843392+00:00"}
{"citation_id": "19930093641", "source_url": "https://ntrs.nasa.gov/api/citations/19930093641/downloads/19930093641.pdf", "page_number": 15, "total_pages": 47, "image_filename": "19930093641_p15.jpg", "text": "13\n\nurements of the sound-pressure level obtained for the three enclosed-engine tractor propellers are of interest. The results obtained at a propeller speed of 3,000 r.p.m. are as follows:\n\n| Tractor position | Sound pressure, decibels |\n|------------------|--------------------------|\n| 1                | 78.5                     |\n| 2                | 78.3                     |\n| 3                | 85.5                     |\n\nThe discrepancy between positions 1 and 2 is probably within the limits of experimental accuracy. In the tunnel tests, the noise level of position 3 corresponded to a roar as compared to a swish for positions 1 and 2. Unfortunately, data were not obtained for the other test arrangements.\n\nPERFORMANCE COMPARISONS\n\nIn order that the merits of the enclosed-engine arrangement may be illustrated, sample performance calculations are presented. The performance of the enclosed-engine arrangement is given only for the case of the pusher arrangement; however, owing to the similarity in the aerodynamic characteristics shown in table I, the computations apply almost equally well to the tractor positions 1 and 2.\n\nHigh speed.— From the measured drag and propulsive efficiencies, the high speeds were computed for four different model conditions (fig. 37). Computations are based on a wing loading of 25.7 pounds per square foot and a power loading of 17.7 pounds per horsepower. The assumed propeller-blade angle of $18-1/2^\\circ$ is lower than the optimum for the high-speed condition, and all the calculated speeds would have been somewhat higher if a larger blade angle had been used. The maximum speeds are as follows:\n\n| Condition                          | High speed m.p.h. |\n|------------------------------------|-------------------|\n| Wing nacelles, tractor:            |                   |\n| 1. With exposed radiators - - - - - | 194               |\n| 2. Without radiators - - - - - - - | 207               |", "timestamp": "2026-07-19T18:24:00.043634+00:00"}