Wide differences of opinion are expressed by automobile builders regarding crankcase-oil dilution. The theories advanced in explanation of dilution fail to elucidate some important facts and must therefore be regarded as unsatisfactory. From a theoretical investigation, the author determines the conditions under which the vapors of various fuels condense during the compression stroke of the engine and, as a result of such analysis, presents the theory that “surface condensation,” or the aggregation of the liquid fuel-particles on the cylinder-walls, is chiefly responsible for crankcase-oil dilution. First, suggested explanations of the dilution are presented, references to previous experiments by several authorities are stated and these are discussed. The effect of jacket-water temperature is analyzed, and whether any condensation of fuel takes place during the compression stroke of a carbureter engine is debated.
THIS report, a discussion of the design problems in heated surface anti-icing equipment, consolidates and compiles all of the heat transfer data known to be available from past experimentation and considered to be required for current and future designs. Consequently, the author believes that the discussion contained herein will be of assistance in some degree to designers and engineers confronted with problems relating to heated surface anti-icing. The report deals with a rapid means of calculating heated wing requirements, charts of heat transfer coefficients, a discussion of instrumentation techniques, and a method of calculating surface temperatures in dry air.
ADEQUATE fuel feeding at altitude, these authors point out, is a matter of vapor elimination, either by preventing its formation or by removing it from the system in the event that its formation cannot be prevented. The effect of vapor is invariably to cause failure of the fuel flow if it forms in sufficient quantity in any part of the fuel system that lies between the fuel tank and the carburetor. This paper gives the results of a study of the conditions that bring about this type of fuel failure, and describes means of exploring the phenomena experimentally so that it can be ascertained in advance of manufacture if remedial steps are necessary. The greatly accelerated rate at which designs of military aircraft with increased performance have been developed, they explain, has added materially to the difficulty of feeding vapor-free fuel to the carburetors at the higher altitudes.
THIS paper presents a simple and rapid method of determining the performance of cross-flow intercoolers, oil coolers, or Prestone radiators from laboratory tests of a model or basic unit of the cooler. The method lends itself equally as well to the determination of the size of a cooler of any set performance. Due to the comparative rapidity with which these calculations can be made, it becomes, with the use of this method, an easy matter to make a series of calculations to determine the relations between lengths, cooling air flow, and pressure drops, for any desired performance.
OIL cooling of aircraft powerplants is increasingly difficult. The weight and drag of the oil coolers necessary with the present maximum “Oil-in” temperature of 185 deg. fahr. (85 deg. cent.) are both decidedly objectionable. It appears possible to increase the “oil-in” temperature to about 220 deg. fahr. (104 deg. cent.) with oils which can be produced by the newer refining methods. The use of an “oil-in” temperature of 220 deg. fahr. would render possible a material reduction in weight, size and drag of oil coolers in comparison with present practice. Oils suitable for use at 220 deg. fahr. “oil-in” temperature would not be likely to cause a material increase of engine-starting difficulty, as they would only be used in summer when the shearing resistance of the oil has slight influence on engine starting. The approximate temperature cycle encountered by the oil in its passage through a modern aircraft-engine is discussed.
INDIRECT or liquid-cooled aircraft engines fit into the picture of future aircraft types better than do the direct or air-cooled engines, the authors contend. As reasons for their belief they draw attention to the small frontal area of this type; the heat capacity of the liquid in equalizing temperatures; and greater freedom in cylinder design because large heat-transfer surfaces are unnecessary. Rolls-Royce has been producing liquid-cooled aero engines for 23 yr, they announce, and has concentrated a large staff on installation problems. One of the results of this work, they report, has been the development of the interchangeable powerplant in which the engine-mounting auxiliaries and bulkhead form a complete detachable unit. These units, the authors explain, are interchangeable within 48 hr, and provide interchangeability between air-cooled and liquid-cooled engines.
ICE formation in the carburetor must depend on, at least, the factors (a) volatility and heat of vaporization of the fuel; (b) mixture ratio; (c) humidity, pressure, and the temperature of the intake air; and (d) heat transfer between the carburetor and its surroundings, especially the engine, according to the authors. Small-scale and full-scale tests were made, descriptions of the seven fuels used and of the testing apparatus being given. The procedures for both sets of tests are outlined and the results are analyzed. Other subjects treated are the heat necessary to melt ice, and correlation with the A.S.T.M. distillation. Five conclusions are stated. Appendix 1 refers to calculation of the relation between intake and mixture temperatures when ice formation occurs. Appendix 2 treats of the construction of equilibrium-air-distillation curves for a series of supplied mixture ratios. Appendix 3 is concerned with engine operation near the danger zone and definition of border conditions.
THIS paper gives the results of an analysis made to determine the proportions of aluminum and steel fins to dissipate maximum quantities of heat for several pressure differences across a finned cylinder. The power required to force the cooling air between the fins and the relative weights of the various designs are presented. The calculation of the heat flow in the fins is based on an experimentally verified, theoretical equation and the surface heat-transfer coefficients and pressure differences were taken from previously reported experiments. In particular, the analysis concerns fin proportions for minimum pressure drop, minimum power, and minimum weight.
The Aircraft Interiors subscription addresses the specialized needs and mechanical requirements for aircraft cabin interior design. The application of these standards will aid in the efficient and cost-effective manufacture of quality aircraft components. The standards in this resource include: Glossary of Technical and Physiological Terms Related to Aerospace Oxygen Systems Flight Deck Layout and Facilities Numeral, Letter and Symbol Dimensions for Aircraft Instrument Displays Oxygen Equipment for Aircraft Human Interface Design Methodology for Integrated Display Symbology
This paper discusses the mechanical, hydraulic, and heat transfer aspects of designing a hydraulic system to gimbal the engine on the Saturn S-IVB Booster Stage. This stage requires restart capability after a moderate period of orbital coast. The design configuration which satisfies the overall functional and environmental requirements is shown. Specific areas of discussion include: 1. Design of a hydraulic pump assembly which must be mounted on an accessory pad which may be as cold as -297°F before pump operation and subjects the pump installation to thermal shock (-297°F to +900°F) during operation. 2. Design of an electric motor driven hydraulic pump for operation in the vacuum of space. The results of certain laboratory tests which have been conducted pursuant to the development of the system are discussed.
With the supersonic airplane's skin hot enough to broil a steak, the cabin atmosphere inside has to be as comfortable and safe as it is in subsonic jet transports. Factors that influence the design of the environmental system, such as amount of fresh air, composition of the cabin atmosphere, heat transfer, penalties, safety, and reliability, are briefly described. From numerous possible systems, examples of three different approaches are given. The importance of establishing a correct and complete problem statement (before expenditure of effort toward solution of the problem) is stressed.
Convective heat transfer coefficients were determined for thermocouple junctions oriented parallel to the gas flow for Reynolds numbers (based on wire diameter) from 163 to 17,500. Chromel-alumel wires of 0.013-0.051 in. diameter were tested in air and products of natural gas combustion at temperatures of 60, 500, and 1000 F. Transient response was used to determine the heat transfer coefficient, and all data were corrected for variation of metal specific heat and radiant heat transfer. The affect on apparent heat transfer coefficient was determined for variations in junction weld-bead size, junction length, and wire separation. An empirical equation has been derived relating Nusselt number and Reynolds number that fits 92% of the test data within ± 10%.
An evaluation of the many new devices proposed, in recent years, for power production. Among these are fuel cells, thermoelectric generators, thermionic generators, and solar cells. Comparisons of these energy converting devices are based on ultimate efficiency (thermodynamic principles), weight, size, and cost, when possible.
Methods of measuring radiation and convective heating on airborne vehicles are discussed. Different modes of heat transfer require special gages for determining transient and steady state heating rates. The slug or slope type calorimeter is used to measure both total and radiant heating. A unique conical purge method is used to insure filter cleanliness during flight duration and measuring period. Various types of heating rate gages are discussed outlining the capabilities of each and the factors influencing their accuracy. Experimental data indicates that convective and conductive heat losses from the sensors can be decreased significantly by incorporating special design features. Black body characteristics of the sensors are discussed in relation to the wave length and temperature of the emitting infrared source. Preflight and postflight calibration techniques are used to eliminate uncertainties.
An approximate analytical method was developed to predict boil-off losses due to external heating of uninsulated missile tanks containing liquid hydrogen. The analysis was applied to the upper-stage tankage of two typical missile configurations which differed only in the type of propulsion utilized; a solid propellant first stage with 90 seconds of action time, and a liquid propellant first stage which burned for 140 seconds. It was predicted that during first-stage flight 15–16% of the liquid hydrogen fuel would be lost due to boil-off. These prohibitive losses could be avoided with a thin layer of light insulation such as corkboard. Therefore, it was concluded that uninsulated missile tanks, containing liquid hydrogen, were not desirable.
Thermodynamic properties of liquid hydrogen and their relationship to the design of rocket engine test facilities are presented. Major points covered are the need to completely define the thermodynamic state of the liquid, and a method of determining the heat gain during flow from the environment and from fluid friction. Charts useful for estimating these effects are included.
The increasing complexity of aircraft structures is creating new problems in structural analysis. The problems are in the areas of static and dynamic loads analysis, stress analysis, statistical methods, fatigue, flutter, acoustics, and associated fields, including aerodynamics and heat transfer. These developments are increasing the need for fully automatic computer programs for design use. Such programs must be technically sound, efficient in the use of computer time, rapid in providing answers for design, and simple enough to minimize training requirements. This paper reviews existing methods from the standpoints of technical requirements, personnel familiarization, and effectiveness. Recommendations are made for future developments.
The need for a comprehensive computing system for aircraft structural analysis is discussed. Suitable criteria for such a system are presented. The Douglas Redundant Force Analysis, a general purpose digital computer method for static stress analysis, is described. Details of the method are presented, including the structure cutter routine for the automatic selection of redundants. Nonlinear and dynamic applications are considered. The FORMAT system described is a comprehensive computing approach that permits rapid programming by the engineer of matrix and pseudo matrix operations. The system provides a means for developing and extending structural analytical programs and for combining such programs with dynamic, aerodynamic, and thermodynamic programs in the solution of structural problems.
Fuel used as a coolant in the supersonic transport can degrade thermally and affect heat transfer surfaces. A heat transfer unit developed by Esso Research to follow the course of degradation reactions and to relate deposit formation to heat transfer rates has been used under a Federal Aviation Agency program to study fuels varying widely in quality. Data reveal that oxygen disappearance, peroxide buildup, and deposit formation are interrelated and can be roughly correlated with fuel composition. Deposit formation above a certain level generally results in loss in heat transfer; but in some cases increases in heat transfer have been observed. The temperature at which significant deposit buildup takes place and loss in heat transfer occurs in the HTU can be predicted by the fuel “breakpoint” measurement made in the ASTM-CRC Fuel Cokers.