Quest for Performance: The Evolution of Modern Aircraft
Chapter 10: Technology of the Jet Airplane
Turbojet and Turbofan Systems
[225] Turbojet and turbofan propulsion systems are employed extensively in jet-powered aircraft. Schematic drawings of the two propulsion systems, taken from reference 133, are given in figure 10.1. The turbojet shown at the top of the figure consists of high- and low-pressure compressors, combustor, and high- and low-pressure turbines. In the turbojet, all the inlet air passes through each element of the engine. The compressors raise the pressure of the inlet air; the pressure ratio varies for different engines but may approach 30 to 1. The high-pressure air enters the combustor where fuel is injected. The fuel-air mixture is ignited and the resulting hot gases pass through the turbines that, in turn, drive the compressors. The exhaust from the turbines provides the thrust that propels the aircraft.
The turbojet shown in figure 10.1(a) is called a twin-spool engine. The low-pressure compressor is driven by the low-pressure turbine,...

illstrated drawing of a turbojet engine
(a) Turbojet


illstrated drawing of a high-bypass-ratio
(b) High-bypass-ratio turbofan
[226Figure 10.1 - Two types of jet propulsion systems.

....and the high-pressure compressor is driven by the high-pressure turbine. These two units rotate at different speeds in order to maintain high efficiency in all stages of compression. The engine shown in figure 10.1(a) has nine stages and seven stages in the low-pressure and high-pressure compressors, respectively, and the low-pressure and high-pressure turbines contain two stages and one stage.
A schematic drawing of a turbofan engine is shown in figure 10.1(b). The turbofan engine contains all the elements of the turbojet shown in figure 10.1(a), but in addition, some of the energy in the hot jet exhaust is extracted by a turbine that drives a fan. A portion of the inlet air that enters the fan is bypassed around the engine; the fan, then, is somewhat like a propeller being driven by the turbomachinery. [227] Unlike the propeller, however, a single fan stage may contain from 20 to 50 blades, is surrounded by a shroud, and is more like a single-stage compressor than a propeller. For example, the pressure ratio across a single fan stage is usually in the range of 1.4 to 1.6; whereas the pressure ratio across the propeller discs of the Lockheed Super Constellation in cruising flight is somewhat less than 1.02.
The bypass ratio of a turbofan engine is defined as the ratio of the mass of air that passes through the fan, but not the gas generator, to that which does pass through the gas generator. Bypass ratios between I and 2 are typical of the first turbofan engines introduced in the early 1960's. The more modern turbofan engines for transport aircraft have bypass ratios that usually fall between 4 and 6, and the engine employed on the Lockheed C-5A has a bypass ratio of 8. The larger the bypass ratio, the greater the amount of energy extracted from the hot exhaust of the gas generator; as much as 75 percent of the total thrust of a turbofan engine may be attributed to the fan.
The single-stage front fan shown on the engine in figure 10.1(b) is integral with the low-pressure compressor, and a four-stage turbine drives both the fan and the compressor. Some turbofan engines are of the three-spool type. The hot gas generator employs two spools, like the turbojet shown in figure 10.1(a), and a third spool that is independent of the other two contains the fan and its own turbines. Fans of more than one stage have also been used, as have aft-mounted fans. The aft-fan design is one in which the fan blades form an extension of an independently mounted turbine situated in the hot exhaust of the gas generator.
Most modern civil and military aircraft are powered by some form of t urbofan engine because such engines consume less fuel to produce a given amount of useful power than do comparable turbojet engines. The higher efficiency of the turbofan engine can be explained with the use of Newton's second law of motion. From this well-known law, it may be deduced that a given level of thrust can be produced at a given flight velocity, either by the addition of a small increment of velocity to a large mass flow of air or by the addition of a large increment of velocity to a small mass flow of air. The required energy addition (fuel), however, is less for the first than for the second case. (A simplified analytical proof of this statement is contained in appendix E.) The improved efficiency of the turbofan as compared with the turbojet is, therefore, directly related to the larger air-flow capacity of the fan engine at a given thrust level.
[228] To give quantitative definition to the preceding discussion, the overall propulsion system efficiency at different speeds is compared in figure 10.2 for several propulsion systems. The overall propulsion system efficiency overall propulsion-system efficiencyis the efficiency with which the energy in the fuel is usefully employed in propelling the aircraft and consists of the product of the cycle efficiency engine cycle efficiency, , and the propulsive efficiency propulsive efficiency, percent, as follows:
overall propulsion-system efficiency expressed as a percent
The cycle efficiency is expressed as the percentage of the heat energy in the fuel that is converted to mechanical energy in the engine, and the....

chart describing the efficiency of a Wright R-3350, Continental 10-520 and an Allison T-56
Figure 10. 2 - Overall efficiency equation of overall efficiency of several types of aircraft propulsion systems. (CP is defined in appendix F)

[229]....propulsive efficiency is expressed as the percentage of' this mechanical energy that IS usefully employed in propelling the aircraft. The methods used in calculating the values of efficiency given in figure 10.2 are described in appendix F.
The Curves in figure 10.2 show the overall propulsion-system efficiency as a function of Mach number for a turbojet and two turbofan engines. The turbojet engine and the turbofan engine of bypass ratio 1.4 have the same gas generator. Both engines show a large increase in efficiency as the Mach number increases. For example, the efficiency of the turbofan with a bypass ratio of 1.4 increases from 8 percent to 27.5 percent as the Mach number is increased from 0.2 to 0.9. The 13-percent improvement in efficiency of the fan engine as compared with the pure jet (at a Mach number of 0.8) results entirely from the addition of the fan. The large increase in efficiency that accompanies an increase in the bypass ratio from 1.4 to 6.0, however, is only partly attributable to the increase in bypass ratio. The overall compressor compression ratio of the engine with bypass ratio of 6.0 is about 25, whereas the corresponding ratio for the other fan engine is about 14. Part of the increase in efficiency shown by the engine of higher bypass ratio is accordingly due to an increase in cycle efficiency.
Also shown in figure 10.2 are points for two reciprocating engines and a turboprop engine. The Wright R-3350 turbocompound engine employed on the Lockheed Super Constellation (see chapter 6) was probably the most efficient reciprocating engine ever designed for aircraft use. The overall efficiency of this engine is shown plotted at the cruising speed of the Constellation. Comparison of the point with the curve for the high bypass ratio turbofan engine indicates that the efficiency of the fan engine is as high as that of the Wright engine at a Mach number twice that at which the Constellation cruised. Thus, the overall propulsion efficiency of the 747 flying at its normal cruising speed is about the same as that of the Constellation at its normal cruising speed. The overall efficiency of the engine with a bypass ratio of 1.4, however, is about 20 percent lower than that of the reciprocating engine even at the normal cruise Mach number of the fan engine of about 0.8. The value of the overall efficiency of the 747 is about 32 percent at a Mach number of 0.8. The trends in figure 10.2 clearly show that, with respect to overall propulsion efficiency, the bypass ratio should increase as the cruising speed decreases, and at some speed the propeller or low-solidarity fan becomes the most efficient type of propulsion system. The selection of the optimum bypass ratio for a [230] particular aircraft, however, involves trade studies of many factors, such as the details of the performance requirements of the aircraft in different flight regimes, the efficiencies of the various components of the engine, and the weight and size of the fan and its installation. Also of importance in the selection of the bypass ratio, particularly for an engine intended for application on a civil aircraft, are the noise characteristics of the engine. Engine noise has not been mentioned so far but Is briefly discussed in a later section of this chapter.
The point indicated by a square symbol in figure 10.2 is for a modern six-cylinder, horizontally opposed, reciprocating engine of the type employed in present-day general aviation aircraft. The value of the efficiency of this engine at a Mach number of 0.3 is about 24 percent as compared with about 17 percent for the turbofan at the same Mach number. The point indicated by a diamond symbol in figure 10.2 is for a contemporary turboprop engine employed in a widely used cargo aircraft. The efficiency of this propulsion system is about the same as that of the turbofan at a Mach number of 0.49. The turboprop engine for which the point is shown in figure 10.2 is an old engine that has a compressor compression ratio of only about 10. An engine of more advanced design would be expected to have a higher value of overall propulsion efficiency. The values of the specific fuel consumption cP, for the reciprocating and turboprop engines were obtained from reference 205.
The preceding paragraphs indicate that the turbofan engine, as compared with the reciprocating engine driving a conventional propeller, offers the following advantages:


  1. The turbofan avoids the compressibility losses that limit the speed at which a propeller-driven aircraft may efficiently cruise.
  2. The weight of the turbofan engine per unit power is significantly less than that of the reciprocating engine.
  3. The turbofan engine is capable of developing a very large amount of power in a single unit without prohibitive mechanical complication.
  4. The overall efficiency of the turbofan propulsion system is about the same as that of the most efficient reciprocating engines ever designed for aircraft use. The turbofan engine attains this efficiency at a higher speed than that which is appropriate for reciprocating engines.
  5. [231] The turbofan engine is more reliable than the reciprocating engine and can be operated many thousands of hours without major maintenance work.  


These five basic reasons explain why the turbojet and turbofan propulsion systems have caused a revolution in aircraft design and in our concepts of the way in which aircraft may be effectively used.

Afterburning Engines
Many military aircraft have need for a large, short-time increase in thrust to be used in such operations as takeoff, climb, acceleration, and combat maneuvers. The afterburning engine provides the answer to this need. In this engine, additional fuel is injected directly into the engine exhaust and burned in the tail pipe. Thrust increases of 50 to 80 percent are achievable by this means in modern engines, but at a large increase in fuel consumption. Afterburner operation is feasible because a jet engine operates at a "lean" fuel-to-air ratio to limit temperatures in the hot, rotating parts of the engine to values consistent with the hightemperature limitations of the materials with which these parts are constructed. Thus, the turbine exhaust contains the excess oxygen necessary for afterburner operation.
Shown in figure 10.3 are sketches (based on ref. 133) of hypothetical turbojet and turbofan engines equipped with afterburners. The rotating elements of these engines are not unlike those of the-nonafterburning types shown in figure 10.1. The long afterburner duct, fuel spray bars, flame holders, and adjustable nozzle distinguish the afterburning engine from its nonafterburning counterpart. Fuel is injected into the exhaust of the rotating part of the engine by the fuel spray bars, and the flame holders stabilize the flame and prevent it from being blown out the end of the tailpipe. To obtain maximum thrust from the engine in both afterburning and nonafterburning operation, an adjustable exhaust nozzle is necessary. A nozzle of continuously varying size and shape would be desirable to maximize performance at all flight and engine-operating conditions. In actual practice, however, a two- or three-position nozzle is usually employed to reduce mechanical complication.
In the turbofan engine shown in figure 10.3(b), afterburning takes place in a mixture of the primary exhaust air and the fan bypass air. In a variation of this design, called the duct-burning turbofan, the spray....

air flow diagram of a turbojet engine
(a) Turbojet
air flow diagram of a turbofan engine
(b) Turbofan
[232Figure 10. 3 - Two types of afterbuming jet propulsion system.

....bars and flame holders are located in the fan duct, and all the afterburning takes place in the bypass air. The bypass ratios employed on afterburning fan engines are usually in the order of 2, much less than is common practice on modern nonafterburning engines for transport aircraft, because afterburning fans are usually found on military aircraft designed to penetrate the transonic and low-supersonic speed ranges. When performance requirements encompass these speed ranges, as well as subsonic flight under various conditions, a low bypass ratio becomes the best compromise.
The afterburner provides a light and mechanically simple means for achieving a large boost in thrust. Fuel consumption with afterburning is large, however, as is engine noise. This latter characteristic is particularly troublesome when afterburning is used for takeoff and initial climb. Afterburning is generally not used in cruising flight except for aircraft and engines specifically designed for long-range supersonic [233] flight. The Anglo-French Concord supersonic transport and the American Lockheed SR-71 supersonic reconnaissance aircraft fall into this latter category.
Thrust Reversers
The amount of force required to stop an aircraft in a given distance after touchdown increases with the gross weight of the aircraft and the square of the landing speed. The size of modern transport aircraft and the speed at which they land makes the use of wheel brakes alone unsatisfactory for routine operations. Most propeller-driven transports produced since World War II employ reversible-pitch propellers to assist in stopping the aircraft on the landing rollout.
The advent of the turbojet and turbofan types of propulsion system required the development of new concepts for augmenting the stopping power provided by the brakes. Some military aircraft deploy one or more parachutes after touchdown as shown in figure 10.4. The aerodynamic drag of the parachutes provides the additional stopping force to augment the brakes. Following each landing, the parachutes must be detached from the aircraft and repacked. The use of these devices for deceleration is not an attractive alternative for any type of routine operations and, by Western standards, is completely unacceptable for commercial airline operations. In contrast, a number of Soviet transport aircraft, including early versions of the Tupelov Tu-134 twinjet transport, were equipped with braking parachutes. Another dragproducing method of assisting aircraft deceleration consists of....

North American XB-70 with landing chutes deployed
Figure 10.4 - North American XB-70 with three drag chutes deployed. [mfr via Martin Copp]


[234]...deploying wing spoilers after the aircraft is on the runway. This technique is routinely used on many Jet-powered transports. (See the section on high-lift systems at the end of this chapter.)

To augment the wheel brakes and aerodynamic drag in decelerating the aircraft, the engines of turbojet- and turbofan-powered transport aircraft are equipped with some form of diverter that, when activated, reverses the thrust and thus provides a powerful stopping force. A schematic drawing of a possible thrust reverser for a high-bypass-ratio turbofan engine is shown in figure 10.5. (See ref. 133.) Both the fan exhaust and the hot exhaust from the gas generator are reversed in the design shown. The elements of the reverser are cascades and clam....

forward thrust mode diagram of jet engine
(a) Reverse thrust configuration
Forward thrust configuration
(b) Forward thrust configuration.
Figure 10.5 - Thrust reverser for turbofan engine.


[235]...shells. A cascade is an array of closely spaced, highly cambered airfoils and is used for changing the direction of airflow; it may also be thought of like the blades of a compressor of constant chord laid out parallel to each other rather than radially about a single axis. The clam shell closes the exhaust nozzle and diverts the gas flow outward and forward.

The engine is shown in the reverse and forward thrust configurations in figures 10.5(a) and 10.5(b), respectively. The fan exhaust is reversed by opening the forward cascade so that the impinging exhaust is turned by the blades in the cascade into the forward direction. In the reverse configuration, the exhaust from the hot gas generator strikes the closed clam shell doors and is diverted forward and outward through circumferential openings in the engine nacelle. Fixed cascades are installed in these openings and aid in turning the exhaust gas forward. In the forward thrust configuration, the stowed clam shell closes the cascade and thus prevents leakage of exhaust gases. The front cascade in the forward thrust configuration is closed on the inside so that the fan exhaust cannot pass through it.
Most thrust reversers employ either or both cascades and clam shells in various configurations depending upon the design of the engine, the bypass ratio, and the type of nacelle in which the engine is mounted. In order to prevent ingestion of hot gases or debris into the engine inlet, the thrust reversers are usually not operated below some minimum speed. This minimum speed depends on the design of the aircraft and engine and their integration; 70 miles per hour is a typical value of the minimum speed for operation of the thrust reverser. Although the primary use of the thrust reverser is to shorten the landing distance, reverse thrust is also employed in flight on some aircraft. In this application, reverse thrust is used when a very rapid, steep descent is required to follow a desired flight profile.
Engine Noise
The preceding paragraphs outline the many advantages of jet propulsion systems. A major disadvantage is the noise problem encountered with these types of propulsion systems applied to commercial transport aircraft. The high noise levels of the propulsion system must be considered in relation to the design of the cabin of the aircraft and to the environment external to the aircraft in the vicinity of the airport. The use of light, effective soundproofing material in the cabin has resulted in interior noise levels that are acceptably low without excessive weight penalty.
[236] The primary impact of the high noise levels associated with jet propulsion systems has been felt by people living in communities surrounding airports from which jet-powered transport aircraft operate. Not only were the early jet transports noisier than contemporary aircraft powered with reciprocating engines, but the increased airline traffic that resulted from the widespread adoption of the jet transport resulted in an increased frequency of aircraft operations at most major airports.
The noise problem became so severe and the associated pressure on the U.S. Congress so great that part 36 of the Federal Air Regulations was formulated and became law on December 1, 1969. These regulations specify certain noise levels that must not be exceeded by new aircraft certified after that date. The regulation further states that all aircraft operated in the United States must comply with the regulations after January 1, 1985.
The present certification process for transport aircraft involves experimental measurements of aircraft noise under controlled conditions. The noise level is measured at specified positions under the approach and climb paths of the aircraft and at a specified position to the side of the runway. The allowable noise levels vary to some extent with the gross weight of the aircraft and thus reflect what is technically possible and realistic. Lower allowable noise levels will no doubt be specified at some future time to reflect advancements in the state of the art.
Aircraft noise reduction has been the subject of intensive research and development for the past two decades. The aircraft and engine manufacturing companies as well as various government research and regulatory organizations have been involved in this work. As a result, much has been learned about methods of noise reduction, and considerable literature exists on the subject.
Four approaches have been followed in the various studies aimed toward reducing aircraft noise. First, much work has been directed toward obtaining an understanding of the basic noise generation and propagation process. Second, new concepts in engine design have been developed to reduce the amount of noise generated at the source. Third, methods for suppressing and absorbing a portion of the noise emanating from the engine have been found. Fourth, aircraft operational techniques have been devised for minimizing noise impact on communities surrounding the airport.
The early jet transports were powered with turbojet-type engines. The hot, high-velocity exhaust is the primary source of noise in this type of propulsion system. The amount of energy in the exhaust that is [237] transformed into noise varies as approximately tile eighth power of the exhaust velocity, and the noise-frequency spectrum is related to the circumference of the exhaust jet. The relative amount of noise energy in the lower frequencies increases as the circumference of the jet increases. Many of the early noise suppressors employed on turbojet propulsion systems were based on the concept of effectively breaking the large exhaust Jet into a number of small jets so that the relative amount of noise at the lower frequencies is reduced. The amount of attenuation that accompanies transmission of the noise through the atmosphere increases as the noise frequencies increase. Thus, by breaking up a large Jet into a number of small jets, the amount of energy transmitted as noise over a given distance is reduced. The noise suppressors shown in figures 10.6(a) and (b) are based on the principle just described.
Another type of noise suppressor proposed for the early turbojetpowered transports is shown in figure 10.6(c). The ejector-type suppressor entrains free-stream air, which is then mixed with the high-velocity exhaust. The velocity of the resulting mixed exhaust is therefore lower than that of the free exhaust of the engine alone, and the noise is accordingly reduced at the source.
A great deal of information has accumulated on the manner in which the various components of the engine should be designed so as to reduce the noise generated by the engine. The turbofan engine and the beneficial effects of increasing the bypass ratio on the propulsive efficiency have been discussed earlier. The advent of the turbofan type of propulsion system had an important effect on the nature of the aircraft noise problem. The extraction of energy from the gas generator for the purpose of driving a fan in a high-bypass-ratio engine would be expected to reduce the noise of the fan engine as compared with a turbojet for the same thrust level. The fan itself, however, was found to constitute a new and highly disturbing source of noise. Studies of the relatively low-bypass-ratio, first-generation fan engines showed that the noise that was propagated from the inlet and the fan discharge ducts was greater than that associated with the high-velocity exhaust from the gas generator.
The noise associated with the fan can be greatly reduced by proper detail design of the fan and by the use of acoustic treatment in certain key areas of the inlet and fan discharge ducts. Acoustic treatment consists in the application of sound absorbing material to the interior passages of the nacelle, as shown in figures 10.7(a) and (b) for short and....

view of corrugated-perimeter-type
(a) Corrugated-perimeter-type noise suppressor.
 view of multiple-tube-type
(b) Multiple-tube-type noise suppressor.
view of ejector-type
(c) Ejector-type noise suppressor.
[238Figure 10.6 - Three types of jet noise suppressor.


....long fan duct installations. (Figure 10.7 was taken from reference 185, which contains a comprehensive summary of basic information dealing with acoustic treatment for noise suppression.) An experimental application of acoustic treatment to the nacelle of a first-generation, low-bypass-ratio turbofan engine is shown in figure 10.8. Most modern high-bypass-ratio engines employ some form of acoustic treatment. The splitter rings shown in figure 10.7 have not been used in any production installations for a number of practical operational reasons, such as possible difficulties in deicing and the possibility of the rings....


acoustic treatment for short  fan duct
(a) Short fan duct.
acoustic treatment for long fan duct
(b) Long fan duct.
[239Figure 10. 7 - Examples of acoustic treatment to short and long fan duct nacelles.(Heavy lines indicate acoustic treatment.)


....being broken by foreign object ingestion with subsequent damage to the rotating parts of the engine.

The development of operational techniques for noise abatement will not be dealt with here other than to indicate that these techniques usually involve (1) selected routing into and out of the airport in order to avoid flight over certain heavily populated areas and (2) the use of power reductions and reduced climb angles on certain segments of the climb following takeoff.
Air Inlets
The tremendous amount of power that can be extracted from a single, modern turbopropulsion system has already been discussed. To generate this power with maximum efficiency, the large quantities of propulsion-system air must be delivered to the engine face with minimum aerodynamic loss, turbulence, and flow distortion. High efficiency must be maintained for different engine-operating conditions, different aircraft speeds and altitudes, and for a wide spectrum of angles of attack and sideslip. As one example, a jet transport must inhale air efficiently in the near static condition at the beginning of takeoff roll, in [240] the relatively low-speed, high-power climb condition, and in high-speed, high-altitude cruise flight. At all flight conditions, the propulsion-system air must be decelerated to a low-speed, high-pressure state at the engine compressor face. The detail design of the air intake and internal flow system determines the efficiency with which the air is delivered to the propulsion system. In this case, the efficiency is defined as the ratio, expressed in percent, of the average total pressure of the air entering the engine to that of the free-stream air. The total pressure is the sum of the static, or ambient, pressure of the air and the impact pressure associated with its motion. Modern jet transports may cruise with values of the pressure recovery, that is efficiency, of 97 to 98 percent. Supersonic aircraft with well-designed, practical inlet and internal flow systems may have pressure recoveries of 85 percent or more for Mach numbers in the 2.0 to 2.5 range.
The demanding requirements for high inlet and internal-flow-system efficiency stimulated a large amount of research, development, and engineering effort in the years following the end of World War II. Fortunately, this effort could be based on a solid foundation of earlier work on such things as cowlings and radiator scoops for piston engines. Inlet activity intensified as aircraft penetrated the transonic and supersonic speed ranges, and the field of inlet and internal flow system design soon became a well-recognized engineering specialty. Especially in modern fighters that may have thrust-to-weight ratios in the order of 1, the inlet and its integration with the airframe exert a powerful influence on the overall aircraft design. The aim in engine-airframe integration is to minimize airplane drag, weight, and complexity and to maximize propulsion-system efficiency while, at the same time, ensuring that the aircraft mission requirements have not been compromised. A detailed discussion of the many facets of inlet design is beyond the scope of the present discussion; however, a few examples of inlets that have been used on civil and military aircraft are illustrated and described in the following paragraphs.
Already shown in the discussion of aircraft noise is an inlet typical of those currently employed on modern subsonic transport and strategic bomber aircraft. The splitter rings in the inlet shown in figure 10.8 are part of an experimental installation, which, as mentioned, are not used on production aircraft. The open nose inlet shown is simple, is light in weight, and when used with a pod-mounted engine, has a short, low-loss duct connecting the engine to the inlet. High-pressure recoveries that are relatively insensitive to normal variations in angle of attack and sideslip are possible with this type of inlet.

Photo of acoustical ring inside of DC8
[241Figure 10. 8 - NASA experimental treated nacelle mounted on McDonnell Douglas DC-8 airplane. [NASA]

In contrast to the pod-type mounting found on so many multiengine transport aircraft, most fighters have one or, at the most, two engines situated inside the fuselage. A variety of inlet locations and designs have been employed to supply air to the propulsion system on these aircraft. Each of these arrangements have both advantages and disadvantages. Four typical installations employed on fighter aircraft are illustrated in figure 10.9. These do not by any means constitute all the successful configurations that have been employed on such aircraft over the years. Most installations, however, are variants of those shown.
The simple nose inlet employed on the North American F-86 fighter is illustrated in figure 10.9(a). As indicated previously, this type of installation enjoys good characteristics through a wide range of angle of attack and sideslip and, when located in the front of the fuselage as contrasted with a pod, is free from aerodynamic interference effects-such as flow separation-from other parts of the aircraft. The long internal duct leading from the inlet to the engine, however, tends to have relatively high pressure losses. In addition, interference between the duct and the pilot's cockpit may be encountered. In some designs, the duct passes under the cockpit; in others, it is split and passes around the cockpit on either side of the pilot. Perhaps the largest drawback of the nose inlet, however, is that neither guns nor radar can be mounted in the front of the fuselage. A nose inlet has not been used on a new fighter in the United States since the early 1950's.
[242] The chin inlet employed on the F-8 airplane shown in figure 10.9(b) has many of the advantages of the simple nose inlet but leaves space in the front of the fuselage for radar or guns and has a somewhat shorter internal duct. Care should be taken in such a design to ensure that at no important flight condition does separated or unsteady flow enter the inlet from the nose of the aircraft. The proximity of the inlet to the ground introduces a possible risk of foreign object ingestion, and, obviously, the nose wheel must be located behind the inlet. The chin inlet, however, is a good choice for some applications and is employed on one new contemporary aircraft (the General Dynamics F-16).
Shown in figure 10.9(c) is the wing-root inlet installation employed on the McDonnell F-101 fighter. Inlets located in this manner offer several advantages. Among these are short, light, internal flow ducts, avoidance of fuselage boundary-layer air ingestion, and freedom to mount guns and radar in the nose of the aircraft. Further, no interference between the cockpit and internal ducting is encountered in this arrangement. The short, curved internal ducts, however, require careful design to avoid flow separation and associated losses, and the inletwing integration must be accomplished in such a way that neither the function of the wing nor the inlet is compromised. Wing-root inlets were used on a number of aircraft in the first decade of the jet fighter, but such inlets are not suitable for modern fighters of high thrust-to-weight ratio because of the large-size inlets required by these aircraft and the difficulty of integrating them with the wing.
Side-mounted inlets as used on the Grumman F11F are illustrated in figure 10.9(d). Used on both single- and twin-engine fighters, the side-mounted inlet arrangement probably offers the best compromise of all the conflicting aerodynamic, structural, weight, and space requirements, and it is used on many modern combat aircraft. Great flexibility in inlet size, shape, vertical position, and fore and aft location is offered by the side-mounted installation. Although the F I IF is a design of the 1950's, side-mounted inlets are used on many fighters of the 1970's and 1980's, as described in chapter 11. Before leaving the discussion of figure 10.9(d), it should be noted that the boundary-layer diverter plates are located so as to prevent ingestion by the inlets of the fuselage boundary-layer air. Such boundary-layer diverters are a feature, really a complication, of all fuselage-mounted inlets.
The inlets just described are of the fixed-geometry type; that is, they do not change shape or size as the aircraft speed varies. Fixed-geometry inlets are suitable for aircraft designed to operate at subsonic...

jet with engine air intake in front center of plane

jet with engine air intake below front center of plane



jet with engine air intake on sides of plane   where wings meet fuselage

jet with engine air intake on sides of plane  before wings meet fuselage and behind cockpit


[243] Figure 10.9 - Four inlet locations used on jet-powered fighter aircraft. [NASA]


[244] and low supersonic speeds. For flight at Mach numbers much beyond 1.6, however, variable-geometry features must be incorporated in the inlet if acceptably high inlet pressure recoveries together with low external drag are to be achieved. This complication is dictated by the physical laws governing the flow of air at supersonic speeds. The nature of supersonic flows is not discussed here, but two variable-geometry inlets are illustrated in figure 10.10. Shown at the top in figure 10.10(a) is the D-type side inlet used on the McDonnell Douglas F-4 fighter. Evident in the photograph are the large fixed diverter plates that also serve to begin compression of the entering flow. The adjustable ramps provide further compression along with the desired variation of throat area with Mach number. The angle of the ramps varies automatically in a prescribed manner as the Mach number changes.

The quarter-round inlet equipped with a translating centerbody or spike, as used on the General Dynamics F-111 airplane, is illustrated in figure 10. 1 0(b). The inlet is seen to be bounded on the top by the wing and on one side by the fuselage. An installation of this type is often referred to as an "armpit" inlet. The spike automatically translates fore and aft as the Mach number changes. Although not evident in the photograph, the throat area of the Inlet also varies with Mach number. This is accomplished by expansion and contraction of the rear part of the spike. The diverter for bypassing the fuselage boundary air is also shown in the photograph. The cover over the inlet is to prevent foreign objects from entering the propulsion system while the aircraft is parked on the ground, and, of course, is removed before flight.
The design of inlet systems for supersonic aircraft is a highly complex matter involving engineering trade-offs between efficiency, complexity, weight, and cost. Some of the factors involved in supersonic inlet design are discussed in references 157 and 179, and the problems of engine-airframe integration on supersonic aircraft are summarized well in reference 180. The highly important problem of selecting and integrating the variable-geometry nozzle of afterburning engines is beyond the scope of the present discussion but is also included in the material presented in reference 180.
Historical Note
This discussion of turbojet and turbofan engines concludes with a few comments on the origins of the propulsion systems. Although rotating turbines and compressors had been in use for various purposes for many years, the idea of coupling the two components, with burners....

jet with adjustable engine air inlet
jet with adjustable engine air inlet></FONT> 
      <FONT FACE=(b)
[245Figure 10. 10 - Two types of variable-geometry inlet. [(a) Arthur L. Shoeni via AAHS; (b) George E. Gillburg via AAHS] between, and utilizing the resultant turbine efflux to propel an aircraft was uniquely that of two men working independently with no knowledge of the other's work. These men were Frank Whittle in England and Hans Joachim Pabst von Ohain in Germany.

Simple as the basic idea was, translation of the turbojet concept into a useful aircraft propulsion system presented formidable problems that required technical innovation and engineering of the highest [246] order. Among the many problems were the design of turbines and compressors of sufficiently high efficiency and the proper matching of these components. If the efficiency of these units was not sufficiently high, the turbine would drive the compressor but have a low velocity exhaust incapable of producing useful thrust. Compressor and turbine efficiencies higher than those of other applications of these components were necessary to produce a usable jet engine.
Any aircraft engine of merit must be light in weight. Satisfaction of this requirement with materials of sufficient strength to withstand, for a protracted length of time, the high-temperature, high-stress environment of the hot rotating parts of the engine was a major problem in early jet engines - and remains with us today as engine temperatures continue to rise in the never-ending quest for increased efficiency. Finally, there remained the myriad detail design problems, such as bearings, lubrication, clearances, methods of fabrication and joining, and so on, that must be solved in any new type of machine. Yet, all these problems were overcome in a rudimentary way in the late 1930's and early 1940's, and useful turbojet engines were first produced in this time period. An exhaustive history of turbojet development, beginning with the water wheel, is given in reference 151. Interesting accounts of the early gestation of the turbojet engine are presented in separate papers by Whittle and von Ohain in reference 140.
Air Commodore Sir Frank Whittle (ret.) is often regarded as the father of modern jet propulsion systems. As a young officer in the British Royal Air Force, he became interested in advanced forms of aircraft propulsion. He tried without success to obtain official support for study and development of his ideas but persisted on his own initiative and received his first patent on jet propulsion in January 1930. With private financial support, he began construction of his first engine in 1935. This engine, which had a single-stage centrifugal compressor coupled to a single-stage turbine, was successfully bench tested in April 1937; it was only a laboratory test rig, never intended for use in an aircraft, but it did demonstrate the feasibility of the turbojet concept.
The firm of Power jets Ltd., with which Whittle was associated, received its first official support in 1938. It received a contract for a Whittle engine, known as the W 1, on July 7, 1939. This engine was intended to power a small experimental aircraft. In February 1940, the Gloster Aircraft Company was chosen to develop the aircraft to be powered by the W1 engine. The vehicle, which would be known today as a research aircraft, was covered by specification E28/39 and is [247] frequently referred to by this designation. It was also known as the Pioneer.
The aircraft that emerged from the Gloster factory in 1941 was z small, single-place, low-wing monoplane equipped with a retractable tricycle landing gear. Air for the engine was supplied by a bifurcate nose inlet that passed the intake air around the pilot in separate duct to the engine located in the rear of the fuselage. The E28/39, which was designed by George Carter of the Gloster Company, weighed 3440 pounds, had a wing span of 29 feet, and was capable of a speed o about 340 miles per hour. The W1 engine installed in the aircraft developed 860 pounds of thrust.
The historic first flight of the Pioneer took place on May 15, 1941, with Flight Lieutenant P. E. G. Sayer as pilot. The aircraft was used for a number of years in the exploration of the problems of flight with jet propulsion and was finally placed in the Science Museum in London in 1946. A brief but interesting account of the development of the E28/39 and its Whittle WI engine, together with a detailed discussion of the first British operational fighter, is given in reference 188.
Great Britain was not the only European nation to show an interest in jet propulsion prior to 1940. The German aircraft manufacturer Ernst Heinkel was searching for new concepts in aircraft propulsion in the mid-1930's. His interest was stimulated when he heard that a young scientist at Goettingen University, Hans Joachim Pabst von Ohain, was investigating a new type of aircraft engine that did not require a propeller. Ohain joined Heinkel in 1936 and continued with the development of his concepts of jet propulsion (ref. 165). A successful bench test of one of his engines was accomplished in September 1937. To avoid the combustor development problems associated with the use of liquid fuel, gaseous hydrogen was employed in this early test demonstration. Later engines used liquid petroleum fuels.
A small aircraft was designed and constructed by Ernst Heinkel to serve as a test bed for the new type of propulsion system. The aircraft, designated the He178, was a shoulder-wing monoplane in which the pilot's enclosed cockpit was placed ahead of the wing and the conventional landing gear (tall-wheel- type) retracted into the side of the fuselage. The air for the 1000-pound thrust engine was supplied by an inlet located in the nose of the fuselage. The fuselage was constructed of metal, and the internally braced wing was made of wood. The wing span of the aircraft was 26 feet, 3 inches; the length was 24 feet, 6 inches; and the area of the wing was 85 square feet. The aircraft weighed about 4000 pounds; and although the maximum speed [248] achieved with the aircraft is not known, the anticipated maximum speed was 527 miles per hour according to reference 201.
The Heinkel He178 flew for the first time on August 27, 1939, almost 2 years before the first flight of the British Gloster E28/39. The pilot on this historic first flight of a jet-powered airplane was Flight Captain Erich Warsitz. Little official interest was shown at this time by the German Government in the new form of propulsion system demonstrated by the He178, and the aircraft was actually flown only a few times before being retired to the Berlin Air Museum. The aircraft was destroyed during an Allied air raid in 1943. Later jet aircraft developments in Germany during World War II are described in references 160 and 201.