ALBERT

Description: Afterburners for turbojet engines have, within the past decade, found increasing application in service aircraft. Practically all engines manufactured today are equipped with some form of afterburner, and its use has increased from what was originally a short-period thrust-augmentation application to an essential feature of the turbojet propulsion system for flight at supersonic speeds. The design of these afterburners has been based on extensive research and development effort in expanded laboratory facilities by both the NACA and the American engine industry. Most of the work of the engine industry, however, has either not been published or is not generally available owing to its proprietary nature. Consequently, the main bulk of research information available for summary and discussion is of NACA origin. However, because industrial afterburner development has closely followed NACA research, the omission is more one of technical detail than method or concept. One principal difficulty encountered in summarizing the work in this field is that sufficient knowledge does not yet exist to rationally or directly integrate the available background of basic combustion principles into combustor design. A further difficulty is that most of the experimental investigations that have been conducted were directed chiefly toward the development of specific afterburners for various engines rather than to the accumulation of systematic data. This work has, nonetheless, provided not only substantial improvements in the performance of afterburners but also a large fund of experimental data and an extensive background of experience in the field. Consequently, it is the purpose of the present chapter to summarize the many, and frequently unrelated, experimental investigations that have been conducted rather than to formulate a set of design rules. In the treatment of this material an effort has been made, however, to convey to the reader the "know how" acquired by research engineers in the course of afterburner studies.

Description: In the early development of jet engines, it was occasionally found that excessive amounts of coke or other carbonaceous deposits were formed in the combustion chamber. Sometimes a considerable amount of smoke was noted in the-exhaust gases. Excessive coke deposits may adversely affect jet-engine performance in several ways. The formation of excessive amounts of coke on or just downstream of a fuel nozzle (figs. 116(a) and (b)) changes the fuel-spray pattern and possibly affects combustor life and performance. Similar effects on performance can result from the deposition of coke on primary-air entry ports (fig. 116(c)). Sea-level or altitude starting may be impaired by the deposition of coke on spark-plug electrodes (fig. 116(b)), deposits either grounding the electrodes completely or causing the spark to occur at positions other than the intended gap. For some time it was thought that large deposits of coke in turbojet combustion chambers (fig. 116(a)) might break away and damage turbine blades; however, experience has indicated that for metal blades this problem is insignificant. (Cermet turbine blades may be damaged by loose coke deposits.) Finally, the deposition of coke may cause high-temperature areas, which promote liner warping and cracking (fig. 116(d)) from excessive temperature gradients and variations in thermal-expansion rates. Smoke in the exhaust gases does not generally impair engine performance but may be undesirable from a tactical or a nuisance standpoint. Appendix B of reference 1 and references 2 to 4 present data obtained from full-scale engines operated on test stands and from flight tests that indicate some effects on performance caused by coke deposits and smoke. Some information about the mechanism of coke formation is given in reference 5 and chapter IX. The data indicate that (1) high-boiling fuel residuals and partly polymerized products may be mixed with a large amount of smoke formed in the gas phase to account for the consistency, structure, and chemical composition of the soft coke in the dome and (2) the hard deposits on the liner are similar to petroleum coke and may result from the liquid-phase thermal cracking of the fuel. During the early development period of jet engines, it was noted that the excessive coke deposits and exhaust smoke were generally obtained when fuel-oil-type fuels were used. Engines using gasoline-type fuels were relatively free from the deposits and smoke. These results indicated that some type of quality control would be needed in fuel specifications. Also noted was the effect of engine operating conditions on coke deposition. It is possible that, even with a clean-burning fuel, an excessive amount of coke could be formed at some operating conditions. In this case, combustor redesign could possibly reduce the coke to a tolerable level. This chapter is a summary of the various coke-deposition and exhaust-smoke problems connected- with the turbojet combustor. Included are (1) the effect of coke deposition on combustor life or durability and performance; (2) the effect of combustor design, operating conditions, inlet variables, and fuel characteristics on coke deposition; (3) elimination of coke deposits; (4) the effect of operating conditions and fuel characteristics on formation of exhaust smoke; and (5) various bench test methods proposed for determining and controlling fuel quality.

Description: Combustion must be maintained in the turbojet-engine combustor over a wide range of operating conditions resulting from variations in required engine thrust, flight altitude, and flight speed. Furthermore, combustion must be efficient in order to provide the maximum aircraft range. Thus, two major performance criteria of the turbojet-engine combustor are (1) operatable range, or combustion limits, and (2) combustion efficiency. Several fundamental requirements for efficient, high-speed combustion are evident from the discussions presented in chapters III to V. The fuel-air ratio and pressure in the burning zone must lie within specific limits of flammability (fig. 111-16(b)) in order to have the mixture ignite and burn satisfactorily. Increases in mixture temperature will favor the flammability characteristics (ch. III). A second requirement in maintaining a stable flame -is that low local flow velocities exist in the combustion zone (ch. VI). Finally, even with these requirements satisfied, a flame needs a certain minimum space in which to release a desired amount of heat, the necessary space increasing with a decrease in pressure (ref. 1). It is apparent, then, that combustor design and operation must provide for (1) proper control of vapor fuel-air ratios in the combustion zone at or near stoichiometric, (2) mixture pressures above the minimum flammability pressures, (3) low flow velocities in the combustion zone, and (4) adequate space for the flame.

Description: From considerations of safety and reliability in performance of gas-turbine aircraft, it is clear that engine starting and acceleration are of utmost importance. For this reason extensive efforts have been devoted to the investigation of the factors involved in the starting and acceleration of engines. In chapter III it is shown that certain basic combustion requirements must be met before ignition can occur; consequently, the design and operation of an engine must be tailored to provide these basic requirements in the combustion zone of the engine, particularly in the vicinity of the ignition source. It is pointed out in chapter III that ignition by electrical discharges is aided by high pressure, high temperature, low gas velocity and turbulence, gaseous fuel-air mixture, proper mixture strength, and-an optimum spark. duration. The simultaneous achievement of all these requirements in an actual turbojet-engine combustor is obviously impossible, yet any attempt to satisfy as many requirements as possible will result in lower ignition energies, lower-weight ignition systems, and greater reliability. These factors together with size and cost considerations determine the acceptability of the final ignition system. It is further shown in chapter III that the problem of wall quenching affects engine starting. For example, the dimensions of the volume to be burned must be larger than the quenching distance at the lowest pressure and the most adverse fuel-air ratio encountered. This fact affects the design of cross-fire tubes between adjacent combustion chambers in a tubular-combustor turbojet engine. Only two chambers in these engines contain spark plugs; therefore, the flame must propagate through small connecting tubes between the chambers. The quenching studies indicate that if the cross-fire tubes are too narrow the flame will not propagate from one chamber to another. In order to better understand the role of the basic factors in actual engine operation, many investigations have been conducted in single combustors from gas-turbine engines and in full-scale engines in altitude tanks and in flight. The purpose of the present chapter is to discuss the results of such studies and, where possible, to interpret these results qualitatively in terms of the basic requirements reported in chapter III. The discussion parallels the three phases of turbojet engine starting: (1) Ignition of the fuel-air mixture (2) Propagation of flame throughout the combustion zone (3) Acceleration of the engine to operating speed.

Description: Studies of the fundamental processes of combustion are usually concerned with wide ranges of investigation of individual processes. In general, each fundamental combustion process may be studied in an environment that is most suited to its evaluation and possibly unrelated basically to any practical application. The majority of the data presented in volume I of this series concern the fundamental aspects of combustion as functions of the individual occurrence of various contributing processes. In a jet engine, however, the various fundamental combustion processes may occur simultaneously and may interact. Furthermore, the engine environment usually does not permit independent variation of single combustion parameters, since specified operating conditions impose specific values on the parameters. In volume II, data are presented to show the effect of operating conditions on the over-all combustion process in different combustion components. To show the effect of operating conditions, it is necessary to specify the range of these conditions within which combustion components may operate. Therefore, this chapter presents only the operating conditions that might be required in the primary combustors and afterburners of typical current turbojet engines. (Corresponding information on ram-jet engines is presented in ch. xisi.) This chapter is not intended to serve as an explanation of engine operation. The operating conditions of the combustion components are presented in terms of total pressures and temperatures at the primary-combustor and afterburner inlets, reference velocities and outlet total temperatures of the primary combustors, and velocities at the plane of the flameholder in the afterburners. The data are presented to relate the operating regions of typical current turbojet combustion components to flight altitudes, Mach numbers, and modes of engine operation. Specifically, data are presented for the combustion parameters of the primary combustor and afterburner of three turbojet engines having rated compressor total-pressure ratios of 5, 8, and 12 under full-throttle conditions. Operational data for the primary combustor also include part-throttle operation at 70, 80, and 90 percent of rated engine speed and windmifling operation. The range of flight conditions includes altitudes from sea level to 65,000 feet and flight Mach numbers from zero to 1.6.

Description: A 1/13-scale model of the forebody of the Republic F-105 with twin-duct wing-root inlets was tested in the Langley 4- by 4-foot supersonic pressure tunnel through a range of angle of attack from -4 deg to 15 deg at a Mach number of 2.01 and a Reynolds number of approximately 3.4 x 10(exp 6) per foot. The tests were made with four configurations which incorporated varying amounts of sweep and stagger of the inlet leading edges, modifications to the areas of the boundary-layer diverter floor plate, and modifications to the area of the boundary-layer diverter bleed slots. The highest overall pressure recovery at an angle of attack of 0 deg (average total-pressure recovery, 0.84 mass-flow ratio, 0.98) was achieved with configuration having an inlet leading-edge sweep angle of 58 deg with no leading-edge stagger. Stagger was found to improve the angle-of- attack performance, but at a sacrifice in inlet efficiency for an angle of attack of 0 deg. The boundary-layer diverter floor height, of the order of one boundary-layer thickness, was satisfactory for bypassing the fuselage boundary layer. The boundary-layer diverter-plate bleed slots were effective in increasing the total-pressure recovery of the inlet. The total-pressure-recovery contour plots, taken at the compressor-face station, indicate the existence of high-velocity "cores" throughout the inlet operating range.

Description: Three highly polished 15- included- angle cone- cylinders with hemispherical tips of several diameters ( 2, 3, and 4 in.) have been flown in order to obtain boundary- layer transition data at very low wall to local stream temperature ratios, and heat- transfer data. All surfaces had a 2-microinch average roughness height. Laminar flow existed over the entire hemispherical nose of the 2- and 3-inch-tip- diameter models throughout the complete flight history. Extreme cooling to wall to local stream temperature ratios at the sonic point as low as 0.20 did not cause transition on the nose for diameters as large as 3 inches. However, extreme cooling did cause early transition on the 4-inch model where it appears probable that transition occurred forward of the 45 station at a wall to local stream temperature ratio of about 0.26. Variations in tip diameter influenced transition downstream of the nose under conditions of extreme cooling. The 2-inch- tip model was laminar at all cone- cylinder stations at temperature ratios as low as 0.32 whereas the 3- and 4-inch-tip models were turbulent at the same local flow conditions but at higher wall to local temperature ratios. Transition on the cone and cylinder of the 3- and 4-inch- tip bodies appeared to be sensitive to local Mach number, and occurred at higher local temperature ratios when values of local Mach number were higher. Increasing the nose diameter from 2 to 3 inches significantly changed the local flow conditions for which laminar flow existed on the cone- cylinder afterbody. However, a further increase in tip size t o a 4-inch diameter had no discernable effect on the local flow conditions at transition. The transition results of the 3- and 4-inch-nose-diameter smooth bodies are similar to those observed on a 7/8-inch-nose-diameter body with roughened surfaces. Turbulent boundary layers resulted in both cases at very low wall to local stream temperature ratios. Both laminar and turbulent heat-transfer data were in good agreement with theoretical Stanton numbers when heat-transfer reduction due to tip blunting was considered.

Description: A design guide is suggested as a basis for indicating combinations of airplane design variables for which the possibilities of pitch-up are minimized for tail-behind-wing and tailless airplane configurations. The guide specifies wing plan forms that would be expected to show increased tail-off stability with increasing lift and plan forms that show decreased tail-off stability with increasing lift. Boundaries indicating tail-behind-wing positions that should be considered along with given tail-off characteristics also are suggested. An investigation of one possible limitation of the guide with respect to the effects of wing-aspect-ratio variations on the contribution to stability of a high tail has been made in the Langley high-speed 7- by 10-foot tunnel through a Mach number range from 0.60 to 0.92. The measured pitching-moment characteristics were found to be consistent with those of the design guide through the lift range for aspect ratios from 3.0 to 2.0. However, a configuration with an aspect ratio of 1.55 failed t o provide the predicted pitch-up warning characterized by sharply increasing stability at the high lifts following the initial stall before pitching up. Thus, it appears that the design guide presented herein might not be applicable when the wing aspect ratios lower than about 2.0.

Description: An investigation of some aspects of the sonic boom has been made with the aid of wind-tunnel measurements of the pressure distributions about bodies of various shapes. The tests were made in the Langley 4- by 4-foot supersonic pressure tunnel at a Mach number of 2.01 and at a Reynolds number per foot of 2.5 x 10(exp 6). Measurements of the pressure field were made at orifices in the surface of a boundary-layer bypass plate. The models which represented both fuselage and wing types of thickness distributions were small enough to allow measurements as far away as 8 body lengths or 64 chords. The results are compared with estimates made using existing theory. To the first order, the boom-producing pressure rise across the bow shock is dependent on the longitudinal development of body area and not on local details. Nonaxisymmetrical shapes may be replaced by equivalent bodies of revolution to obtain satisfactory theoretical estimates of the far-field pressures.

Description: Time histories of noise pressures near ground level were measured during flight tests of fighter-type airplanes over fairly flat, partly wooded terrain in the e Mach number range between 1.13 and 1.4 and at altitudes from 25,000 to 45,000 feet. Atmospheric soundings and radar tracking studies were made for correlation with the measured noise data. The measured and calculated values of the pressure rise across the shock wave were generally in good agreement. There is a tendency for the theory to overestimate the pressure at locations remote from the track and to underestimate the pressures for conditions of high tailwind at altitude. The measured values of ground-reflection factor averaged about 1.8 f or the surface tested as compared to a theoretical value of 2.0. P o booms were measured in all cases. The observers also generally reported two booms; although, in some cases, only one boom was reported. The shock-wave noise associated with some of the flight tests was judged to be objectionable by ground observers, and in one case the cracking of a plate-glass store window was correlated in time with the passage of the airplane at an altitude of 25,000 feet.