TURBOJET ENGINES

The turbojet engine has found widespread use in aircraft propulsion because of the relatively high power output per powerplant weight and size. Very few aircraft powerplants can com­pare with the high output, flexibility, simplic­ity, and small size of the aircraft gas turbine. The coupling of the propeller and recipro­cating engine is one of the most efficient means

known for converting fuel energy into propul­sive energy. However, the intermittent action of the reciprocating engine places practical limits to the airflow that can be processed and restricts the development of power. The con­tinuous, steady flow feature of the gas turbine allows such a powerplant to process consider­ably greater airflow and, thus, utilize a greater expenditure of fuel energy. While the pro­pulsive efficiency of the turbojet engine is con­siderably below that of the reciprocating en­gine-propeller combination, the specific power output of the turbojet at high speeds is quite superior.

The operation of the turbojet engine involves a relatively large change in velocity being im­parted to the mass flow through the engine. Figure 2.6 illustrates the operation of a typical turbojet engine by considering the processing given a unit weight of inlet airflow. Consider a unit weight of ambient air approaching the inlet to the engine then experiencing the changes in pressure and volume as it is proc­essed by’the turbojet. The chart of pressure versus volume of figure 2.6 shows that the unit weight of airflow at atmospheric condition A is delivered to the inlet entrance at condition B. The purpose of the inlet or diffuser j^s to reduce the velocity and increase the pressure of the flow entering the compressor section. Thus, the aerodynamic compression produces an increase in pressure and decrease in volume of the unit weight of air and delivers air to the compressor at cond ition C. The work done by the aerodynamic compression of the inlet or diffuser is represented by the area ABCX. Generally, most conventional turbojet engines require that the compressor inlet flow be sub­sonic and supersonic flight will involve con­siderable aerodynamic compression in the inlet.

Air delivered to the compressor inlet at con­dition C is then subject to further compression through the compressor section. As a result of the function of the compressor, the unit weight of air is subject to a decrease in volume and increase in pressure to condition D. The compressor pressure ratio should be high to produce a high thermal efficiency in the engine The area XCDZ represents the work done by the compressor during the compression of the unit weight of air. Of course, certain losses and inefficiencies are incurred during the com­pression and the power required to operate the compressor will be greater than that indicated by the work done on the engine airflow.

Compressed air is discharged from the com­pressor to the combustion chamber at condition D. Fuel is added in the combustion chamber and the combustion of fuel liberates consider­able heat energy. The combustion process in the gas turbine differs from that of the recipro­cating engine in that the process is essentially a constant pressure addition of heat energy. As a result, the combustion of fuel causes a large change in temperature and large change of volume of the unit weight of airflow. The process in the combustion chamber is repre­sented by the change from point D to point E of the pressure-volume diagram of figure 2.6.

The combustion products are delivered to the turbine section where sufficient work must be extracted to power the compressor section. The combustion chamber discharges high tem­perature, high pressure gas to the turbine where a partial expansion is accomplished with a drop in pressure and increase in volume to point F on the pressure-volume diagram. The work extracted from the unit weight of air by the turbine section is represented by the area ZEFY. As with the compressor, the actual shaft work extracted by the turbine will differ from that indicated by the pressure-volume diagram because of certain losses incurred through the turbine section. For steady, sta­bilized operation of the turbojet engine the power extracted by the turbine will equal the power required to operate the compressor. If the turbine power exceeds the compressor power required, the engine will accelerate; if the turbine power is less than the compressor power required, the engine will decelerate.

The partial expansion of the gases through the turbine will provide the power to operate the engine. As. the gases are discharged from the turbine at point F, expansion will continue through the tailpipe nozzle until atmospheric pressure is achieved in the exhaust. Thus, continued expansion in the jet nozzle will re­duce the pressure and increase the volume of the unit weight of air to point G on the pressure volume diagram. As a result, the final jet velocity is greater than the inlet velocity and the momentum change necessary for the de­velopment of thrust has4 been created. The area YFGA represents the work remaining to provide the expansion to jet velocity after the turbine has extracted the work required to operate the compressor.

Of course, the combustion chamber discharge could be more completely expanded through a larger turbine section and the net power could be used to operate a propeller rather than pro­vide high exhaust gas velocity. For certain applications, the gas turbine-propeller combi­nation could utilize the high power capability of the gas turbine with greater propulsive efficiency.

FUNCTION OF THE COMPONENTS. Each of the engine components previously de­scribed will contribute some function affecting the efficiency and output of the turbojet engine. For this reason, each of these components should be analyzed to determine the require­ments for satisfactory operating characteristics.

The inlet or diffuser must be matched to the powerplant to provide the compressor entry with the required airflow. Generally, the compressor inlet must receive the required air­flow at subsonic velocity with uniform dis­tribution of velocity and direction at the compressor face. The diffuser must capture high energy air and deliver it at low Mach number uniformly to the compressor. When the inlet is along the sides of the fuselage, the edges of the inlet must be located such that the inlet receives only high energy air and provision must be made to dispose of the boundary layer along the fuselage surface. At supersonic flight speeds, the diffuser must slow the air to subsonic with the least waste of energy in the inlet air and accomplish the process with a minimum of aerodynamic drag. In addition, the inlet must be efficient and stable in operation throughout the range of angles of attack and Mach numbers of which the airplane is capable.

The operation of the compressor can be af­fected greatly by the uniformity of flow at the compressor face. When large variations in flow velocity and direction exist at the face of the axial compressor, the efficiency and stall – surge limits are lowered. Thus, the flight conditions which involve high angle of attack and high sideslip can cause deterioration of inlet performance.

The compressor section is one of the most im­portant components of the turbojet engine. The compressor must furnish the combustion chamber with large quantities of high pressure air in a most efficient manner. Since the com­pressor of a jet engine has no direct cooling, the compression process takes place with a minimum of heat loss of the compressed air. Any friction loss or inefficiency of the com­pression process is manifested as an undesirable additional increase in the temperature of the compressor discharge air. Hence, compressor efficiency will determine the compressor power necessary to create the pressure rise of a given airflow and will affect the temperature change which can take place in the combustion chamber.

The compressor section of a jet engine may be an axial flow or centrifugal flow compressor. The centrifugal flow compressor has great util­ity, simplicity, and flexibility of operation. The operation of the centrifugal compressor requires relatively low inlet velocities and a plenum chamber or expansion space must be provided for the inlet. The impeller rotating at high speed receives the inlet air and pro­vides high acceleration by virtue of centrifugal force. As a result, the air leaves the impeller

at very high velocity and high kinetic energy. A pressure rise is produced by subsequent ex­pansion in the diffuser manifold by converting the kinetic energy into static pressure energy. The manifold then distributes the high pres­sure discharge to the combustion chambers. A double entry impeller allows a given diam­eter compressor to process a greater airflow. The major components of the centrifugal com­pressor are illustrated in figure 2.7.

The centrifugal compressor can provide a relatively high pressure ratio per stage but the provision of more than one or two stages is rarely feasible for aircraft turbine engines. The single stage centrifugal compressor is capable of producing pressure ratios of about three or four with reasonable efficiency. Pres­sure ratios greater than four require such high impeller tip speed that compressor efficiency decreases very rapidly. Since high pressure ratios are necessary to achieve low fuel con­sumption, the centrifugal compressor finds greatest application to the smaller engines where simplicity and flexibility of operation are the principal requirements rather than high efficiency.

The axial flow compressor consists of alter­nate rows of rotating and stationary airfoils. The major components of the axial flow com­pressor are illustrated in figure 2.7. A pressure rise occurs through the row of rotating blades since the airfoils cause a decrease in velocity relative to the blades. Additional pressure rise takes place through the row of stationary blades since these airfoils cause a decrease in the absolute velocity of flow. The decrease I in velocity, relative or absolute, effects a com – | pression of the flow and causes the increase in static pressure. While the pressure rise per stage of the axial compressor is relatively low, the efficiency is very high and high pressure ratios can be obtained efficiently by successive axial stages. Of course, the efficient pressure rise in each stage is limited by excessive gas velocities. The multistage axial flow com­pressor is capable of providing pressure ratios from five to ten (or greater) with efficiencies which cannot be approached with a multi­stage centrifugal compressor.

The axial flow compressor can provide efficiently the high pressure ratios necessary for low fuel consumption. Also, the axial compressor is capable of providing high air­flow with a minimum of compressor diameter. When compared with the centrifugal com­pressor, the design and construction of the axial compressor is relatively complex and costly and the high efficiency is sustained over a much narrower range of operating conditions. For these reasons, the axial compressor finds greatest application where the demands of efficiency and output predominate over con­siderations’ of cost, simplicity, flexibility of operation, etc. Multi spool compressors and variable stator blades serve to improve the operating characteristics of the axial com­pressor and increase the flexibility of operation.

The combustion chamber must convert the fuel chemical energy into heat energy and cause a large increase in the total energy of the engine airflow. The combustion chamber will oper­ate with one principal limitation: the dis­charge from the combustion chamber must be at temperatures which can be tolerated by the turbine section. The combustion of liquid hydrocarbon fuels can produce gas temperatures which are in excess of 1,700 to 1,800° C. However, the maximum continuous turbine blade operating temperatures rarely exceed 800° to 1,000° C and considerable excess air must be used in the combustion chamber to prevent exceeding these temperature limits.

. While the combustion chamber design may take various forms and configurations, the main features of a typical combustion chamber are illustrated by figure 2.8. The combustion chamber receives the high pressure discharge from the compressor and introduces approxi­mately one half of this air into the immediate area of the fuel spray. This primary combus­tion air must be introduced with relatively high turbulence and quite low velocities to

NAVWEPS 00-80T-80

airplane performance

maintain, a nucleus of combustion in the com­bustion chamber. In the normal combustion process, the speed of flame propagation is quite low and, if the local velocities are too high at the forward end of the combustion chamber, poor combustion will result and it is likely that the flame will blow out. The secondary air—or cooling flow—is introduced downstream from the combustion nucleus to dilute the com­bustion products and lower the discharge gas temperature.

The fuel nozzie must provide a finely – atomized, evenly distributed spray of fuel through a wide range of flow rates. Very specialized design is necessary to provide a nozzle with suitable characteristics. The spray pattern and circulation in the combustion chamber must make efficient use of the fuel by complete combustion. The temperatures in the combustion nucleus can exceed 1,700° to 1,800° C but the secondary air will dilute the gas and reduce the temperature to some value which can be tolerated in the turbine section. A pressure drop will occur through the com­bustion chamber to accelerate the combustion gas rearward. In addition, turbulence and fluid friction will cause a pressure drop but this loss must be held to the minimum incurred by providing complete combustion. Heat trans­ferred through the walls of the combustion chamber constitutes a loss of thermal energy and should be held to a minimum. Thus, the combustion chamber should enclose the com­bustion space with a minimum of surface area to minimize heat and friction losses. Hence, the ‘‘annular” typ; combustion chamber offers certain advantages over the multiple “can” type combustion chamber.

The turbine section is the most critical element of the turbojet engine. The function of the turbine is to extract energy from the combus­tion gases and furnish power to drive the com­pressor and accessories. In the case of the turboprop engine, the turbine section must ex­tract a very large portion of the exhaust gas energy to drive the propeller in addition to the compressor and accessories.

The combustion chamber delivers high en­ergy combustion gases to the turbine section at high pressure and tolerable temperature. The turbine nozzle vanes are a row of stationary blades immediately ahead of the rotating tur­bine. These blades form the nozzles which discharge the combustion gases as high ve­locity jets onto the rotating turbine. In this manner, the high pressure energy of the com­bustion gases is converted into kinetic energy and a pressure and temperature drop takes place. The function of the turbine blades operating in these jets is to develop a tangen­tial force along the turbine wheel thus extract­ing mechanical energy from the combustion gases. This is illustrated in figure 2.8.

The form of the turbine blades may be a com­bination of two distinct types. The impulse type turbine relies upon the nozzle vanes to accomplish the conversion of combustion gas static pressure to high velocity jets. The impulse turbine blades are shaped to produce a large deflection of the gas and develop the tangential force by the flow direction change. In such a design, negligible velocity and pres­sure drop occurs with the flow across the tur­bine rotor blades. The reaction type turbine differs in that large velocity and pressure changes occur across the turbine rotor blades. In the reaction turbine, the stationary nozzle vanes serve only to guide the combustion gas onto the turbine rotor with negligible changes in velocity and pressure. The reaction tur­bine rotor blades are shaped to provide a pres­sure drop and velocity increase across the blades and the reaction from this velocity in­crease provides the tangential force on the wheel. Generally, the turbine design is a form utilizing some feature of each of the two types.

The turbine blade is subjected to high centrifugal stresses which vary as the square of the rotative speed. In addition, the blade is subjected to the bending and torsion of the tangential impulse-reaction forces. The blade must withstand these stresses which are generally of a vibratory and cyclic nature while at high temperatures. The elevated temperatures at which the turbine must func­tion produce extreme conditions for struc­tural creep and fatigue considerations. Conse­quently, the engine speed and temperature op­erating limits demand very careful considera­tion. Excessive engine temperatures or speeds may produce damage which is immediately apparent. However, creep and fatigue damage is cumulative and even though damage may not be immediately apparent by visual inspec­tion, proper inspection methods (other than visual) must be utilized and proper records kept regarding the occurrence.

Actually, the development of high tempera­ture alloys for turbines is a critical factor in the development of high efficiency, high output aircraft gas turbines. The higher the tem­perature of gases entering the turbine, the higher can be the temperature and pressure of the gases at discharge from the turbine with greater exhaust jet velocity and thrust.

The function of the tailpipe or exhaust notice is to discharge the exhaust gases to the atmos­phere at the highest possible velocity to pro­duce the greatest momentum change and thrust. If a majority of the expansion occurs through the turbine section, there remains only to con­duct the exhaust gases rearward with a mini­mum. energy loss. However, if the turbine operates against a noticeable back pressure, the nozzle must convert the remaining pressure energy into exhaust gas velocity. Under ideal conditions, the nozzle would expand the flow to the ambient static pressure at the exhaust and the area distribution in the nozzle must provide these conditions. When, the ratio of exhaust gas pressure to ambient pressure is relatively low and incapable of producing sonic flow, a converging nozzle provides the expan­sion. The exit area must be of proper size to bring about proper exit conditions. If the exit

area is too large, incomplete expansion will take place; if the exit area is too small, an over expansion tendency results. The exit area can affect the upstream conditions and must be properly proportioned for overall performance.

When the ratio of exhaust gas pressure to ambient pressure is greater than some critical value, sonic flow can exist and the nozzle will be choked or limited to some maximum flow. When supersonic exhaust gas velocities are re­quired to produce the necessary momentum change, the expansion process will require the convergent-divergent nozzle illustrated in fig­ure 2.9. With sufficient pressure available the initial expansion in the converging portion is subsonic increasing to sonic velocity at the. throat. Subsequent expansion in the divergent portion of the nozzle is supersonic and the re­sult is the highest exit velocity for a given pressure ratio and mass flow. When the pres­sure ratio is very high the final exit diameter required to expand to ambient pressure may be very large but is practically limited to the fuselage or nacelle afterbody diameter. If the exhaust gases exceed sonic velocity, as is possi­ble in a ramjet combustion chamber or after­burner section, only the divergent portion of the nozzle may be necessary.

Figure 2.9 provides illustration of the func­tion of the various engine components and the changes in static pressure, temperature, and velocity through the engine. The conditions at the inlet provide the initial properties of the engine airflow. The compressor section fur­nishes the compression pressure rise with a certain unavoidable but undesirable increase in temperature. High pressure air delivered to combustion chamber receives heat from the combustion of fuel and experiences a rise in temperature. The fuel flow is limited so that the turbine inlet temperature is within limits which can be tolerated by the turbine structure. The combustion takes place at relatively con­stant pressure and initially low velocity. Heat addition then causes large increases in gas vol­ume and flow velocity.

NOZZLE TYPES

CONVERGENT NOZZLE CONVERGENT-DIVERGENT NOZZLE

ENGINE OPERATING CONDITIONS

Generally, the overall fuel-air ratio of the turbojet is quite low because of the limiting turbine inlet temperature. The overall air – fuel ratio is usually some value between 80 to 40 during ordinary operating conditions be­cause of the large amount of secondary air or cooling flow.

High temperature, high energy combustion gas is delivered to the turbine section where power is extracted to operate the compressor section. Partial or near-complete expansion can take place through the turbine section with the accompanying pressure and temperature drop. The exhaust nozzle completes the ex­pansion by producing the final jet velocity and momentum change necessary in the develop­ment of thrust.

TURBOJET OPERATING CHARACTER­ISTICS. The turbojet engine has many oper­ating characteristics wh’ich are of great im­portance to the various items of jet airplane performance. Certain of these operating char­acteristics will provide a strong influence on the range, endurance, etc., of the jet-powered airplane. Other operating characteristics will require operating techniques which differ greatly from more conventional powerplants.

The turbojet engine is essentially a thrust – producing powerplant and the propulsive power produced is a result of the flight speed. The variation of available thrust with speed is relatively small and the engine output is very nearly constant with flight speed. The mo­mentum change given the engine airflow de­velops thrust by the following relationship:

Ta=Q(Va-Vd

where

Та — thrust available, lbs. ig=mass flow, slugs per sec.

F^inlet or flight velocity, ft. per sec.

Fa=jet velocity, ft. per see.

Since an increase in flight speed will increase the magnitude of Vu a constant thrust will be obtained only if there is an increase in mass flow, jQ, or jet velocity, V2, When at low velocity, an increase in velocity will reduce the velocity change through the engine with­out a corresponding increase in mass flow and the available thrust will decrease. At higher velocity, the beneficial ram helps to overcome this effect and the available thrust no longer decreases, but increases with speed.

The propulsive power available from the turbojet engine is the product of available thrust and velocity. The propulsive horse­power available from the turbojet engine is related by the following expression: where

Pa=propulsive power available, h. p.

*T* .L – ____ ‘ 1 _ 1 1 11

i#=uuusi av ana Die, ids.

V = flight velocity, knots

The factor of 325 evolves from the use of the nautical unit of velocity and implies that each pound of thrust developed at 325 knots is the equivalent of one horsepower of propul­sive power. Since the thrust of the turbojet engine is essentially constant with speed, the power available increases almost linearly with speed. In this sense, a turbojet with 5000 lbs. of thrust available could produce a propulsive power of 5,000 h. p. at 325 knots or 10,000 h. p. at 650 knots. The tremendous propulsive power at high velocities is one of the principal features of the turbojet engine. When the engine RPM and operating altitude arc fixed, the variation with speed of turbojet thrust and power available is typified by the first graph of figure 2.10.

The variation of thrust output with engine speed is a factor of great importance in the operation of the turbojet engine. By reason­ing that static pressure changes depend on the square of the flow velocity, the changes of pressure throughout the turbojet engine would

be expected to vary as the square of the rota­tive speed, N. However, since a variation in rotative speed will alter airflow, fuel flow, compressor and turbine efficiency, etc., the thrust variation will be much greater than just the second power of rotative speed. In­stead of thrust being proportional to N2, the typical fixed geometry engine develops thrust approximately proportional to N3-6. Of course, such a variation is particular to constant alti­tude and speed.

Figure 2.10 illustrates the variation of per­cent maximum thrust with percent maximum RPM for a typical fixed geometry engine. Typical values from this graph are as follows:

Percent max. thrust

100 (of course)

96.5

83.6

69.2 45-8

28.7

Note that in the top end of power output, each 1 percent RPM change causes a 3- 5-percent change in thrust output. This illustrates the power of variation of thrust with rotative speed which, in this example, is N3‘6. Also note that the top 20 percent of RPM controls more than half of the output thrust.

While the fixed geometry engine develops thrust approximately proportional to N33, the engine with variable geometry will demonstrate a much more powerful effect of rotative speed. When the jet engine is equipped with a vari­able nozzle, multispool compressor, variable stator blades, etc., the engine is more likely to develop thrust proportional to rotative speed from values of N4"5 to N60. For ex­ample, if a variable geometry engine develops thrust proportional to N50, each one per cent RPM change causes a 5-0-percent thrust change at the top end of power output. Also, the top 13 percent of RPM would control the top 50 percent of thrust output.

The powerful variation of thrust with engine speed has certain ramifications which should
be appreciated. If the turbojet powerplant operates at less than the “trimmed” or adjusted speed for maximum thrust, the deficiency of thrust for takeoff may cause a considerable increase in takeoff distance. Du-ring approach, an excessively low RPM may cause very low thrust and produce a very steep glide path. In addition, the low RPM range involves the much greater engine acceleration time to pro­duce thrust for a waveoff. Another compli­cation exists when the thrust is proportional to some large power of rotative speed, e. g., N®°. The small changes in RPM produce such large variations in thrust that instruments other than the tachometer must be furnished for accurate indication of thrust output.

The “specific fuel consumption, c” is an important factor for evaluating the perform­ance and efficiency of operation of a turbojet engine. The specific fuel consumption is the proportion between the fuel flow (in lbs. per hr.) and the thrust (in lbs.). For example, an engine which has a fuel flow of 14,000 lbs. per hr. and a thrust of 12,500 lbs. has a specific fuel consumption of:

Fuel flow

c‘~ Thrust

14,0 lbs./hr.

Ct~ 12,500 lbs.

c,= 1.12 lbs./hr./lb.

Thus, each unit pound of thrust requires 1.12 lbs. per hr. fuel flow. Obviously, high engine efficiency would be indicated by a low value of et. Typical values for turbojet engines with relatively high pressure ratios range from 0.8 to 1.2 at design operating conditions in sub­sonic flight. High energy fuels and greater pressure ratios tend to produce the lower values of ct. Supersonic flight with the attendant in­let losses and high compressor inlet air tem­peratures tend to increase the specific fuel con­sumption to values of 1.2 to 2.0. Of course, the use of an afterburner is quite inefficient

О 10 20 30 40 50 60 70 80 90 100
PERCENT MAXIMUM RPM

due to the low combustion pressure and values of c, from 2.0 to 4.0 are typical with after­burner operation.

The turbojet engine usually has a strong preference for high RPM to produce low specif­ic fuel consumption. Since the normal rated thrust condition is a particular design point for the engine, the minimum value of c, will occur at or near this range of RPM. The illustration of figure 2.10 shows a typical vari­ation of c, with percent maximum RPM where values of RPM less than 80 to 85 percent pro­duce a specific fuel consumption much greater than the minimum obtainable. This pref­erence for high RPM to obtain low values of c, is very pronounced in. the fixed geometry engine. Turbojet engines with multispool compressors tend to be less sensitive in this respect and are more flexible in their operating characteristics. Whenever low values of c, are necessary to obtain range or endurance, the preference of the turbojet engine for the design operating RPM can be a factor of great influence.

Altitude is one factor which strongly affects the performance of the turbojet engine. An increase in altitude produces a decrease in density and pressure and, if below the tropo – pause, a decrease in temperature. If a typical nonafterburning turbojet engine is operated at a constant RPM and true airspeed, the varia­tion of thrust and specific fuel consumption with altitude can be approximated from figure 2.11. The variation of density in the standard atmosphere is shown by the values of density ratio at various altitudes. Typical values of the density ratio at specific altitudes are as follows:

Altitude, ft.: Sea level

5,000.

10,000

22,000

55.000

40.000

50,000

If the fixed geometry engine is operated at a constant V (TAS) in subsonic flight and con­stant N (RPM) the inlet velocity, inlet ram, and compressor pressure ratio are essentially constant with altitude. An increase in alti­tude then causes the engine air mass flow to decrease in a manner very nearly identical to the altitude density ratio. Of course, this de­crease in mass flow will produce a significant effect on the output thrust of the engine. Actually, the variation of thrust with altitude is not quite as severe as the density variation because favorable decreases in temperature occur. The decrease in inlet air temperature will provide a relatively greater combustion gas energy and allow a greater jet velocity. The increase in jet velocity somewhat offsets the decrease in mass flow. Of course, an in­crease in altitude provides lower temperatures below the tropopause. Above the tropopause, no further favorable decrease in temperature takes place so a more rapid variation of thrust will take place. The approximate variation of thrust with altitude is represented by figure 2.11 and some typical values at specific alti­tudes are as follows :

Thrust at altit]uU ,Tkrust at sta tml}

……….. 1.000

……………….. 888

…………………. 785

…………………. 604

…………………. 392

…………………. 315

……………….. 180

Since the change in density with altitude is quite rapid at low altitude turbojet takeoff per­formance will be greatly affected at high alti­tude. Also note that the thrust at 35,000 ft. is approximately 39 percent of the sea level value.

The thrust added by the afterburner of a turbojet engine is not affected so greatly by altitude as the basic engine thrust. The use of afterburner may provide a thrust increase of 50 percent at low altitude or as much as 100 per­cent at high altitude.

(QUANTITY) AT SEA LEVEL Figure 2.11. Approximate Effect of Altitude on Engine Performance

When the inlet ram and compressor pressure ratio is fixed, the principal factor affecting the specific fuel consumption is the inlet air temp­erature, When the inlet air temperature is lowered, a given heat addition can provide relatively greater changes in pressure or vol­ume. As a result, a given thrust output requires less fuel flow and the specific fuel con­sumption, c„ is reduced. While the effect of altitude on specific fuel consumption does not compare with the effect on thrust output, the variation is large enough to strongly influence range and endurance conditions. Figure 2.11 illustrates a typical variation of specific fuel consumption with altitude. Generally, the specific fuel consumption decreases steadily with altitude until the tropopause is reached and the specific fuel consumption at this point is approximately 80 percent of the sea level value.

Above the tropopause the temperature is con­stant and altitudes slightly above the tropo­pause cause no further decrease in specific fuel consumption. Actually, altitudes much above the tropopause bring about a general deteriora­tion of overall engine efficiency and the specific fuel consumption begins an increase with altitude. The extreme altitudes above the tropopause produce low combustion chamber pressures, low compressor Reynolds Numbers, low fuel flow, etc. which are not conducive to high engine efficiency.

Because of the variation of c, with altitude, the majority of turbojet engines achieve maxi­mum efficiency at or above 35,000 ft. For this reason, the turbojet airplane will find optimum range and endurance conditions at. or above 35,000 ft. provided the aircraft is not thrust or compressibility limited at these altitudes.

The governing apparatus of the turbojet engine consists primarily of the items which control the flow of fuel to the engine. In addition, there may be included certain functions which operate variable nozzles, variable stator vanes, variable inlets, etc. Generally, the fuel Con­trol and associated items should regulate fuel flow, nozzle area, etc. to provide engine per­formance scheduled by the throttle or power lever. These regulatory functions provided must account for variations in altitude, tem­perature, and flight velocity.

One principal governing factor which must be available is that a selected power setting (RPM) must be maintained throughout a wide range of flight conditions. Figure 2.12 illus­trates the variation of fuel flow with RPM for a turbojet operating at a particular set of flight conditions. Curve 1 depicts the varia­tion with RPM of the fuel flow required for stabilized, steady state operation of the engine. Each point along this curve 1 defines the fuel flow which is necessary to achieve equilib­rium at a given RPM. The steady state fuel flow produces a turbine power to equal the compressor power requirement at a particular RPM. The throttle position primarily com­mands a given engine speed and, as changes occur in the ambient pressure, temperature, and flight speed, the steady state fuel flow will vary. The governing apparatus must account for these variations in flight conditions and maintain the power setting scheduled by throttle position.

In addition to the maintenance of steady state operation, the fuel control and associ­ated engine control items must provide for the transient conditions of engine acceleration and deceleration, In order to accelerate the en­gine, the fuel control must supply a fuel flow greater than that required for steady state operation to produce a turbine power greater than the compressor power requirement. How­ever, the additional fuel flow to accelerate the engine must be controlled and regulated to prevent any one or combination of the follow­ing items:

(1) compressor stall or surge

(2) excessive turbine inlet temperature

(3) excessively rich fuel-air ratio which

may not sustain combustion Generally, the stall-surge and turbine tem­perature limits predominate to form an ac­celeration fuel flow boundary typified by curve

ALL CURVES APPROPRIATE FOR A PARTICULAR:

2 of figure 2.12. Curve 2 of this illustration defines an upper limit of fuel flow which can be tolerated within stall-surge and tempera­ture limits. The governing apparatus of the engine must limit the acceleration fuel flow within this boundary.

To appreciate the governing requirements during the acceleration process, assume the engine described in figure 2.12 is in steady state stabilized operation at point A and it is desired to accelerate the engine to maximum RPM and stabilize at point C. As the throttle is placed at the position for maximum RPM, the fuel control will increase the fuel flow to point В to provide acceleration fuel flow. As the engine accelerates and increases RPM, the fuel control will continue to increase the fuel flow within the acceleration boundary until the engine speed approaches the controlled maxi­mum RPM at point C. As the engine speed nears the maximum at point C, the fuel control will reduce fuel flow to produce stabilized oper­ation at this point and prevent the engine overspeeding the commanded RPM. Of course, if the throttle is opened very gradually, the acceleration fuel flow is barely above the steady state condition and the engine does not ap­proach the acceleration fuel flow boundary. While this technique is recommended for ordinary conditions to achieve trouble free operation and good service life, the engine must be capable of good acceleration to produce rapid thrust changes for satisfactory flight control.

In order for the powerplant to achieve mini­mum acceleration times, the fuel control must provide acceleration fuel flow as close as practical to the acceleration boundary. Thus, a maximum controlled acceleration may pro­duce limiting turbine inlet temperatures or slight incipient stall-surge of the compressor. Proper maintenance and adjustment of the engine governing apparatus is essential to produce minimum acceleration times without incurring excessive temperatures or heavy stall – surge conditions.

During deceleration conditions, the mini­mum allowable fuel flow is defined by the lean limit to support combustion. If the fuel flow is reduced below some critical value at each RPM, lean blowout or flameout will occur. This condition is illustrated by curve 3 of figure 2.12 which forms the deceleration fuel flow boundary. The governing apparatus must regulate the deceleration fuel flow within this boundary.

To appreciate the governing requirements during the deceleration process, assume the engine described in figure 2.12 is in stabilized, steady state operation at point C and it is desired to decelerate to idle conditions and stabilize at point E. As the throttle is placed at the position for idle RPM, the fuel control will decrease the fuel flow to point D to provide the deceleration fuel flow. As the engine decelerates and decreases RPM, the fuel gov­erning will continue to decrease the fuel flow within the deceleration boundary until the idle fuel flow is reached and RPM is established at point E. Of course, if the throttle is closed very slowly, the deceleration fuel flow is barely below the steady state condition and the engine does not approach the deceleration fuel flow boundary. The fuel control must provide a deceleration flow close to the boundary to provide rapid decrease in thrust and satisfactory flight control.

In most cases, the deceleration fuel flow boundary is considerably below the steady state fuel flow and no great problem exists in obtaining satisfactory deceleration character­istics. In fact, the greater problem is con­cerned with obtaining proper acceleration characteristics. For the majority of centrifu­gal flow engines, the acceleration boundary is set usually by temperature limiting conditions rather than compressor surge conditions. Pea к operating efficiency of the centrifugal com­pressor is obtained at flow conditions which are below the surge limit, hence acceleration fuel flow boundary is determined by turbine temperature limits. The usual result is that

the centrifugal flow engine has relatively large acceleration margins and good acceleration characteristics result with the low rotational inertia. The axial flow compressor must oper­ate relatively close to the stall-surge limit to obtain peak efficiency. Thus, the acceleration fuel flow boundary for the axial flow engine is set by these stall-surge limits which are more immediate to steady state conditions than tur­bine temperature limits. The fixed geometry axial flow engine encounters relatively small acceleration margins and, when compared to the centrifugal flow engine with larger accel­eration margins and lower rotational inertia, has inferior acceleration characteristics. Cer­tain variation of the axial flow engine such as variable nozzles, variable stator blades, multi­ple-spool compressors, etc., greatly improve the acceleration characteristics.

A note of caution is appropriate at this point. If the main fuel control and govern­ing apparatus should malfunction or become inoperative and an unmodulated secondary or emergency system be substitued, extreme care must be taken to avoid abrupt changes in throttle position. In such a case, very gradual movement of the throttle is necessary to ac­complish changes in power setting without excessive turbine temperatures, compressor stall or surge, or flameout.

There are various instruments to relate ІШт portant items of turbojet engine performance. Certain combinations of these instruments are capable of immediately relating the thrust output of the powerplant in a qualitative man­ner. It is difficult to provide an instrument or combination of instruments which immedi­ately relate the thrust output in a quantitative manner. As a result, the pilot must rely on a combination of instrument readings and judge the output performance according to standard values particular to the powerplant. Some of the usual engine indicating instruments are as follows:

(1) The tachometer provides indication of

engine speed, N, by percent of the maximum

RPM. Since the variation of thrust with RPM is quite powerful, the tachometer in­dication is a powerful reference.

(2) The exhaust gas temperature gauge provides an important reference for engine operating limitations. While the tempera­ture probe may be located downstream from the turbine (tailpipe or turbine discharge temperature) the instrument should provide an accurate reflection of temperatures up­stream in the turbine section. The exhaust gas temperature relates the energy change accomplished by fuel addition.

(3) The fuel flowmeter can provide a fair reflection of thrust output and operating efficiency. Operation at high density alti­tude or high inlet air temperatures – reduces the output thrust and this effect is related by a reduction of fuel flow.

(4) The tailpipe total pressure (jp’+q in the tailpipe) can be correlated with the jet thrust for a given engine geometry and set of operating conditions. The: output thrust can be related accurately with various com­binations of compressor inlet total pressure, tailpipe total pressure, ambient pressure and temperature. Hence, pressure differential (Af), pressure ratio, and tailpipe total pres­sure instruments can provide more accurate immediate indications of output thrust than combined indications of RPM and EGT. This is especially true with variable geom­etry or multiple spool engines.

Many other specialized instruments furnish additional information for more detailed items of engine performance. Various additional engine information is realized from fuel pres­sure, nozzle positions, compressor inlet air temperature, etc.

TURBOJET OPERATING LIMITATIONS.

The operating characteristics of the turbojet engine provide various operating limitations which must be given due respect. Operation of the powerplant within the specified limita­tions is absolutely necessary in order to obtain the design service life with trouble-free opera­tion. The following items describe the critical areas encountered during the operational use of the turbojet engine:

(1) The limiting exhaust gas temperatures pro­vide the most important restrictions to the op­eration of the turbojet engine. The turbine components are subject to centrifugal loads of rotation, impulse and reaction loads on the blades, and various vibratory loads which may be inherent with the design. When the turbine components are subject to this variety of stress in the presence of high temperature, two types of structural phenomena must be considered. When a part is subject to a certain stress at some high temperature, creep failure will take place after a period of time. Of course, an increase in temperature or stress will increase the rate at which creep damage is accumulated and reduce the time required to cause failure. An­other problem results when a part is subjected to a repeated or cyclic stress. Fatigue failure will occur after a number of cycles of a varying stress. An increase in temperature or magni­tude of cyclic stress will increase the rate of fatigue damage and reduce the number of cycles necessary to produce failure. It is important to note that both fatigue and creep damage are cumulative.

A gross overstress or overtemperature of the turbine section will produce damage that is immediately apparent. However, the creep and fatigue damage accumulated through pe­riods of less extreme’ overstress or overtem­perature is more subtle. If the turbine is subject to repeated excessive temperatures, the greatly increased rate of creep and fatigue damage will produce failure early within the anticipated service life.

Generally, the operations which produce the highest exhaust gas temperatures are starting, acceleration, and maximum thrust at high altitude. The time spent at these temperatures must be limited arbitrarily to prevent excessive accumulation of creep and fatigue. Any time spent at temperatures in excess of the operational limits for these con­ditions will increase the possibility of early failure of the turbine components.

While the turbine components are the most critically stressed high temperature elements they are not the only items. The combustion chamber components may be critical at low altitude where high combustion chamber pres­sures exist. Also, the airframe structure and equipment adjacent to the engine may be sub­ject to quite high temperatures and require provision to prevent damage by excess time at high temperature.

(2) The compressor stall or surge has the pos­sibility of producing damaging temperatures in the turbine and combustion chamber or un­usual transient loads in the compressor. While the stall-surge phenomenon is possible with the centrifugal compressor, the more common. occurrence is with the axial flow compressor. Figure 2.13 depicts the pressure distribution that may exist for steady state operation of the engine. In order to accelerate the engine to a greater speed, more fuel must be added to increase the turbine power above that required to operate the compressor.

Suppose that the fuel flow is increased be­yond the steady state requirement without a change in rotative speed. The increased com­bustion chamber pressure due to the greater fuel flow requires that the compressor dis­charge pressure be higher. For the instant before an engine speed change occurs, an in­crease in compressor discharge pressure will be accompanied by a decrease in compressor flow velocity. The equivalent effect is illustrated by the flow components onto the rotating com­pressor blade of figure 2.13. One component of velocity is due to rotation and this compo­nent remains unchanged for a given rotative velocity of the single blade. The axial flow velocity for steady state operation combines with rotational component to define a result­ant velocity and direction. If the axial flow component is reduced, the resultant velocity and direction provide an increase in angle of

COMPRESSOR STALL

COMPRESSOR COMBUSTION TURBINE EXHAUST

CHAMBER NOZZLE

Figure 2,13. Effect of Compressor Stall and Inlet Temperature on Engine Operation

attack for the rotating blade with a subsequent increase in pressure rise. Of course, if the change in angle of attack or pressure rise is beyond some critical value, stall will occur. While the stall phenomenon of a series of rotating compressor blades differs from that of a single airfoil section in a free airstream, the cause and effect are essentially the same.

If an excessive pressure rise is required through the compressor, stall may occur with the attendant breakdown of stable, steady flow through the compressor. As stall occurs, the pressure rise drops and the compressor does not furnish discharge at a pressure equal to the combustion chamber pressure. As a result, a flow reversal or backfire takes place. If the stall is transient and intermittent, the indica­tion will be the intermittent "bang” as back­fire and flow reversal take place. If the stall develops and becomes steady, strong vibration and a loud (and possibly expensive) roar develops from the continuous flow reversal. The increase in compressor power required tends to reduce RPM and the reduced airflow and increased fuel flow cause rapid, immediate rise in exhaust gas temperature. The pos­sibility of damage is immediate with the steady stall and recovery must be accomplished quickly by reducing throttle setting, lowering the airplane angle of attack, and increasing airspeed. Generally, the compressor stall is caused by one or a combination of the fol­lowing items:

(a) A malfunctioning fuel control or gov­erning apparatus is a common cause. Proper maintenance and adjustment is a necessity for stall-free operation. The malfunctioning is most usually apparent during engine acceleration.

(b) Poor inlet conditions are typical at high angles of attack and sideslip. These conditions reduce inlet airflow and create nonuniform flow conditions at the com­pressor face. Of course, these conditions are at the immediate control of the pilot.

(r) Very high altitude flight produces low compressor Reynolds numbers and an effect similar to that of airfoil sections. As a decrease to low Reynolds numbers reduces the section clmai very high altitudes reduce the maximum pressure ratio of the com­pressor. The reduced stall margins increase the likelihood of compressor stall.

Thus, the recovery from a compressor stall must entail reduction of throttle setting to reduce fuel flow, lowering angle of attack and sideslip and increasing airspeed to improve inlet condition, and reducing altitude if high altitude is a contributing factor.

(3) While the flameout is a rare occurrence with modern engines, various malfunctions and operating conditions allow the flameout to remain a possibility. A uniform mixture of fuel and air will sustain combustion within a relatively wide range of fuel-air ratios. Com­bustion can be sustained with a fuel-air ratio as rich as one to five or as lean as one to twenty – five. Fuel air ratios outside these limits will not support combustion due to the deficiency of air or deficiency of fuel. The characteristics of the fuel nozzle and spray pattern as well as the governing apoaratus must insure that the nucleus of combi, .ion is maintained through­out the range of engine operation.

If the rich limit of fuel-air ratio is exceeded in the combustion chamber, the flame will blow out. While this condition is a pos­sibility the more usual cause of a flameout is exceeding the lean blowout limit. Any con­dition which produces some fuel-air ratio leaner than the lean limit of combustion will produce a flameout. Any interruption of the fuel supply could bring on this condition. Fuel system failure, fuel system icing, or pro­longed unusual attitudes could starve the flow of fuel to the engine. It should be noted the majority of aviation fuels are capable of holding in solution a certain small amount of water. If the aircraft is refueled with rela­tively warm fuel then flown to high altitude,

the lower temperatures can precipitate this water out of solution in liquid or ice crystal form.

High altitude flight produces relatively small air mass flow through the engine and the rela­tively low fuel flow rate. At these conditions a malfunction of the fuel control and governing apparatus could cause flameout. If the fuel control allows excessively low fuel flow during controlled deceleration, the lean blow out limit may be exceeded. Also, if the governed idle condition allows any deceleration below the idle condition the engine will usually continue to lose speed and flameout.

Restarting the engine in flight requires suffi­cient RPM and airflow to allow stabilized op­eration. Generally, the extremes of altitude are most critical for attempted airstart.

(43 An increased compressor inlet air tempera­ture can have a profound effect on the output thrust of я turbojet engine. As shown in figure 2.13, an increase in compressor inlet temperature produces an even greater increase in the compressor discharge temperature. Since the turbine inlet temperature is limited to some maximum value, any increase in com­pressor discharge temperature will reduce the temperature change which can take place in the combustion chamber. Hence, the fuel flow will be limited and a reduction in thrust is incurred.

The effect of inlet air temperature on thrust output has two special ramifications. At take­off, a high ambient air temperature at a given pressure altitude relates a high density altitude. Thus, the takeoff thrust is reduced because of low density and low mass flow. In addition to the loss of thrust due to reduced mass flow, thrust and fuel flow are reduced further be­cause of the high compressor inlet temperature. In flight at high Mach number, the aerodynamic heating will provide an increase in compressor inlet temperature. Since the compressor inlet temperature will reflect the compressor dis­charge temperature and the allowable fuel flow, the compressor inlet air temperature may provide a convenient limit to sustained high speed flight.

(5} The effect of engine overspeed or critical vi­bration speed ranges is important in the service life of an engine. One of the principal sources of turbine loads is the centrifugal loads due to rotation. Since the centrifugal loads vary as the square of the rotative speed, a 5 percent overspeed would produce 10.25 percent over­stress (1.05* = 1-10253- The large increase in stress with rotative speed could produce very rapid accumulation of creep and fatigue dam­age at high temperature. Repeated overspeed and, hence, overstress can cause failure early in the anticipated service life.

Since the turbojet engine is composed of many different distributed masses and elastic structure, there are certain vibratory modes and frequencies for the shaft, blades, etc. While it is necessary to prevent any resonant conditions from existing within the normal operating range, there may be certain vibra­tory modes encountered in the low power range common to ground operation, low altitude endurance, acceleration or deceleration. If certain operating RPM range restrictions are specified due to vibratory conditions, opera­tions must be conducted with a minimum of time in this area. The greatly increased stresses common to vibratory conditions are quite likely to cause fatigue failures of the offending components.

The operating limitations of the engine are usually specified by various combinations of RPM, exhaust gas temperature, and allowable time. The conditions of high power output and acceleration have relatively short times allowable to prevent abuse of the powerplant and obtain good service life. While the al­lowable times at various high power and acceleration condition appear arbitrary, the purpose is to reduce the spectrum of loading which contributes the most rapid accumulation of creep and fatigue damage. In fact, in some instances, the arbitrary time standards can be set to suit the particular requirements of a certain type of operation. Of course, the effect on service life of any particular load spectrum must be anticipated.

One exception to the arbitrary time standard for operation at high temperatures or sus­tained high powers is the case of the after­burner operation. When the cooling flow is only that necessary to prevent excessive tem­peratures for adjacent structure and equipment, sustained operation past a time limit may cause damage to these items.

THRUST AUGMENTATION. Many op­erating performance conditions may require that additional thrust be provided for short periods of time. Any means of augmenting the thrust of the turbojet engine must be ac­complished without an increase in engine speed or maximum turbine section temperature. The various forms of afterburning or water injection allow the use of additional fuel to provide thrust augmentation without increase in engine speed or turbine temperature.

The afterburner is a relatively simple means of thrust augmentation and the principal fea­tures are light weight and large thrust increase. A typical afterburner installation may add only 10 to 20 percent of the basic engine weight but can provide a 40- to 60-percent increase in the static sea level thrust. The afterburner con­sists of an additional combustion area aft of the turbine section with an arrangement of fuel nozzles and flameholders. Because the local flow velocities in the afterburner are quite high, the flameholders are necessary to provide the turbulence to maintain combustion within the afterburner section. The turbojet engine operates with airflows greatly in excess of that chemically required to support combus­tion of engine fuel. This is necessary because of cooling requirements and turbine tempera­ture limitations. Since only 15 to 30 percent of the engine airflow is used in the combustion chamber, the large excess air in the turbine discharge can support combustion of large amounts of additional fuel. Also, there are no highly stressed, rotating members in the afterburner and very high temperatures can be tolerated. The combustion of fuel in the after­burner brings additional increase in tempera­ture and volume and’ adds considerable energy to the exhaust-gases producing increased jet velocity. The majoT components of the after­burner are illustrated in figure 2.14.

One necessary feature of the turbojet engine equipped with afterburner is a variable nozzle area. As the afterburner begins functioning, the exit nozzle area must increase to accom­modate the increased combustion products. If the afterburner were to begin functioning without an increase in exit area, the mass flow through the engine would drop and the tem- peratnres would increase rapidly. The nozzle area must be controlled to increase as after­burner combustion-begins. As a result, the engine mass flow is given a large increase in jet velocity with the corresponding increase in thrust. ■ –

The combustion of fuel in the afterburner takes place at low pressures and is relatively inefficient. This basic inefficiency of the low pressure combustion is given evidence by the large increase in specific fuel combustion. Generally, the use of afterburner at least will double the specific fuel consumption. As an example, consider a turbojet engine capable of producing 10,000 lbs. of thrust which can develop 15,000 lbs.-of thrust with the use of afterburner. Typical values for specific fuel consumption would-be r< = 1.05 for the basic engine or r( = 2.1 when the afterburner is in use. The fuel flow during operation would be as follows:

fuel flow = (thrust) (specific fuel consump­tion)

without afterburner,

fuel flow = (10,000) (1.05)

= 10,500 lbs./hr.

with afterburner,

fuel flow = (15,000) (2.1)

= 31,500 lbs./hr.

The low efficiency of the afterburner is illus­trated by the additional 21,000 lbs./hr. of fuel flow to create the additional 5,000 lbs. of

AFTERBURNER COMPONENTS

WATER INJECTION

Figure 2.14. Thrust Augmentation and the Gas Turbine-Propeller Combination

thrust. Because of the high fuel consumption during afterburner operation and the adverse effect on endurance, the use of the afterburner should be limited to short periods of time. In addition, there may be limited time for the use of the afterburner due to critical heating of supporting or adjacent structure in the vicin­ity of the afterburner.

The specific fuel consumption of the basic engine will increase with the addition of the afterburner apparatus. The losses incurred by the greater fluid friction, nozzle and flame – holder pressure drop, etc. increase the specific fuel consumption of the basic engine approxi­mately 5 to 10 percent.

The principal advantage of afterburner is the ability to add large amounts of thrust with relatively small weight penalty. The applica­tion of the afterburner is most common to the interceptor, fighter, and high speed type aircraft. .

The use of water injection in the turbojet en­gine is another means of thrust augmentation which allows the combustion of additional fuel within engine speed and temperature limits. The most usual addition of water injection de­vices is to supplement takeoff and climbout performance, especially at high ambient tem­peratures and high altitudes. The typical water injection device can produce a 25 to 35 percent increase in thrust.

The most usual means of water injection is direct flow of the fluid into the combustion chamber. This is illustrated in figure 2.14. The addition of the fluid directly into the com­bustion chamber increases the mass flow and reduces the turbine inlet temperature. The drop in temperature reduces the turbine power and a greater fuel flow is required to maintain engine speed. Thus, the mass flow is increased, more fuel flow is allowed within turbine limits, and greater energy is imparted to the exhaust gases.

The fluid injected into the combustion cham­bers is generally a mixture of water and alco­hol. The water-alcohol solution has one immediate advantage in that it prevents fouling of the plumbing from the freezing of residual fluid at low temperatures. In addition, a large concentration of alcohol in the mixture can provide part of the additional chemical energy required to maintain engine speed. In fact, the large concentration of alcohol in the in­jection mixture is a preferred means of adding additional fuel energy. If the added chemical energy is included with the water flow, no abrupt changes in governed fuel flow are necessary and there is less chance of underspeed with fluid injection and overspeed or over­temperature when fluid flow is exhausted. Of course, strict proportions of the mixture are necessary. Since most water injection devices are essentially an unmodulated flow, the use of this device is limited to high engine speed and low altitude to prevent the water flow from quenching combustion.

THE GAS TURBINE-PROPELLER COM­BINATION. The turbojet engine utilizes the turbine to extract sufficient power to operate the compressor. The remaining exhaust gas energy is utilized to provide the high exhaust gas velocity and jet thrust. The propulsive efficiency of the turbojet engine is relatively low because thrust is produced by creating a large velocity change with a relatively small mass flow. The gas turbine-propeller combin­ation is capable of producing higher propulsive efficiency in subsonic flight by having the pro­peller operate on a much greater mass flow.

The turboprop or prop jet powerplant re­quires additional turbine stages to continue expansion in the turbine section and extract a very large percent of the exhaust gas energy as shaft power. In this sense, the turboprop is primarily a power producing machine and the jet thrust is a small amount of the output propulsive power. Ordinarily, the jet thrust of the turboprop accounts for 15 to 25 percent of the total thrust output. Since the turbo­prop is primarily a power producing machine,

the turboprop powerplant is rated by an ‘‘equivalent shaft horsepower.”

eshp^bhpa-Щ-

325i?„

where

ESHP=equivalent shaft horsepower BHP= brake horsepower, or shaft horse­power applied to the propeller Ту = jet thrust, lbs.

V=flight velocity, knots, TAS Up = propeller efficiency

The gas turbine engine is capable of processing large quantities of air and can produce high output power for a given engine size. Thus, the principal advantage of the turboprop powerplant is the high specific power output, high power per engine weight and high power per engine size.

The gas turbine engine must operate at quite high rotative speed to process large airflows and produce high power. However, high rotative speeds are not conducive to high propeller efficiency because of compressibility effects. A large reduction of shaft speed must be provided in order to match the powerplant and the propeller. The reduction gearing must provide a propeller shaft speed which can be utilized effectively by the propeller and, be­cause of the high rotative speeds of the turbine, gearing ratios of 6 to 15 may be typical. The transmission of large shaft horsepower with such high gearing involves considerable design problems to provide good service life. The problems of such gearing were one of the greatest difficulties in the development of turboprop powerplants.

The governing apparatus for the turboprop powerplant must account for one additional variable, the propeller blade angle. If the propeller is governed separately from the tur­bine, an interaction can exist between the engine and propeller governers and various “hunting,” overspeed, and overtemperature conditions are possible. For this reason, the engine-propeller combination is operated at a constant RPM throughout the major range of output power and the principal variables of con­trol are fuel flow and propeller blade angle. In the major range of power output, the throttle commands a certain fuel flow and the propeller blade angle adjusts to increase the propeller load and remain at the governed speed.

The operating limitations of the turboprop powerplant are quite similar in nature to the operating limitations of the turbojet engine. Generally, the turbine temperature limitations are the most critical items. In addition, over­speed conditions can produce overstress of the gearing and propeller as well as overstress of the turbine section.

The performance of the turboprop illustrates the typical advantages of the propeller-engine combination. Higher propulsive efficiency and high thrust and low speeds provide the characteristic of range, endurance, and takeoff performance superior to the turbojet. As is typical of all propeller equipped powerplants, the power available is nearly constant with speed. Because the power from the jet thrust depends on velocity, the power available in­creases slightly with speed. However, the thrust available decreases with speed. The equivalent shaft horsepower, ESHP, of the turboprop is affected by mass flow and inlet temperature in fashion similar to that of the turbojet. Thus, the ESHP will vary with altitude much like the thrust output of the turbojet because the higher altitude produces much lower density and engine mass flow. The gas turbine-propeller combination utilizes a number of turbine stages to extract shaft power from the exhaust gases and, as high compressor inlet temperatures reduce the fuel flow allowable within turbine temperature limits, hot days will cause a noticeable loss of output power. Generally, the turboprop is just as sensitive, if not more sensitive, to com­pressor inlet air temperature as the turbojet engine.

The specific fuel consumption of the turbo­prop powerplant is defined as follows:

specific fuel consumption =

_____ engine fuel flow

equivalent shaft horsepower

__ lbs. per hr.

ESHP

Typical values for specific fuel consumption, c, range from 0.5 to 0.8 lbs. per hr. per ESHP. The variation of specific fuel consumption with operating conditions is similar to that of the turbojet engine. The minimum specific fuel consumption is obtained at relatively high power setting and high altitudes. The low inlet air temperature reduces the specific fuel consumption and the lowest values of c are ob­tained near altitudes of 25,000 to 35,000 ft. Thus, the turboprop as well as the turbojet has a preference for high altitude operation.