Category AIRCRAF DESIGN

Turbo props are described in Section 10.4.4. They are very similar to turbojets and turbofans except that the high energy of the exhaust jet is utilized to drive a pro­peller by incorporating additional low-pressure turbine stages, as shown in Fig­ure 10.7. Thrust developed by the propellers is the propulsive force for the aircraft. A small amount of residual thrust could be left at the nozzle exit plane, which should be added to the propeller thrust. The relationship between the thrust power (TP) and the gas turbine SHP is related to propeller efficiency, nprop, as:

TP = SHP X Vop + F X 1% (10.20)

ESHP is a convenient way to define the combination of shaft and jet power, as follows:

ESHP = TP/Vop = SHP + (F x VoVnprop (10.21)

Aircraft at a static condition have an ESHP = SHP because the small thrust at the exit nozzle is not utilized. As speed increases, ESHP > SHP, as there is some thrust at the nozzle. SFC and specific power are expressed in terms of ESHP.

Summary

The formulae provide good reasoning for the gas turbine domain of application, as shown in Figure 10.2. Turboprops provide the best economy for a design flight speed at and below Mach 0.5 and are well suited for shorter ranges of operation. At higher speeds, up to Mach 0.98, turbofans with a high BPR provide better efficiencies (see the comments following Equation 10.5). At supersonic speeds, the BPR is reduced and, in most cases, uses an AB. Smaller aircraft have piston engines up to a certain size (i. e., « <500 HP). Above 500 HP, turboprops prove better than piston engines.

Low pressure High pressure

Compressor Compressor (shaded)

(ECS)

Figure 10.14. Installation effects

10.2 Engine Integration with an Aircraft: Installation Effects

Engine manufacturers typically supply bare engines to aircraft manufacturers, which install them to integrate with an aircraft. The same type of engines can be used by different aircraft manufacturers; each has its own integration requirements. Installing an engine in an aircraft is a specialized technology with which aircraft designers must be knowledgeable. Engine integration is accomplished by aircraft manufacturers in consultation with engine manufacturers.

A bare engine at the test stand performs differently than an installed engine on an aircraft. The installation effects of an engine result from having a nacelle – that is, the losses of intake and exhaust plus off-takes of power (e. g., driving motors and generators) and air-bleeds (e. g., anti-icing and environmental control). The total loss of thrust at takeoff could be as high as 8 to 10% of what is generated by a bare engine at the test bed; at cruise, the loss can be reduced to less than 5%. Figure 10.14 shows typical off-takes that are required due to various installation effects. Designers conduct analytical and empirical studies to establish key parame­ters in order to arrive at a design that produces satisfactory thrust to meet aircraft – performance requirements.

A nacelle is the housing for the engine and it interfaces with an aircraft; typ­ically, it is pod/pylon-mounted in civil aircraft designs. A nacelle on an aircraft with more than one engine is pod/pylon-mounted on the wing and/or the fuselage. Propeller-driven engine nacelles are also similar to podded nacelles, modified by the presence of a propeller (see Section 10.7.2). An aircraft with one engine is aligned in the plane of the aircraft symmetry; engines with propellers can have a small lateral inclination of 1 or 2 deg about the aircraft centerline to counter the slip­stream and gyroscopic effects from a rotating propeller. As discussed previously, wing-mounted nacelles are best for relieving wing-bending in the flight load. The engines on military aircraft are buried in the fuselage and therefore do not have a nacelle unless the designer chooses to have pods (e. g., some older designs). Military – aircraft designers must consider intake design as described in Section 10.8.2. The

position of the nacelle relating to the aircraft and the shaping to reduce drag are important considerations (see Section 9.8).

Figure 10.13 is a schematic diagram showing the station numbers for an AB jet engine. To keep numbers consistent with the turbojet numbering system, there is no difference between Stations 4 and 5, which represent the turbine exit condition. Station 5 is the start and Station 6 is the end of AB. Station 7 is the final exit plane. Figure 10.13 also shows the isentropic AB cycle in a T-s diagram.

AB is deployed only in military aircraft (except in the civil supersonic Con­corde) as a temporary thrust-augmentation device to meet the mission demand at takeoff and/or fast acceleration and maneuvers to engage or disengage in combat. AB is applied at full throttle by activating a fuel switch. The pilot can feel the deploy­ment by the sudden increase in the g-level in the flight direction. A ground observer

notices a sudden increase in the noise level, which can exceed the physical thresh­old. An AB glow is visible at the exit nozzle; in the dark, it appears as a spectac­ular plume with supersonic expansion “diamonds.” In the absence of any down­stream rotating machines, the AB temperature limit can be increased from 2,000 to 2,200 deg K, at the expense of a significant increase in the fuel flow (i. e., a richer fuel-to-air ratio than in the core combustion).

An AB exit nozzle invariably runs choked and requires a convergent-divergent nozzle for the supersonic expansion to increase the gain in momentum for the thrust augmentation. Typically, to gain a 50% thrust increase, fuel consumption increases from 100 to 120%; that is why it is used only for a short period, not necessarily in one burst. It is interesting that AB in bypass engines is an attractive proposition because the AB inlet temperature is lower. In fact, all modern combat-category engines use a low bypass of 1 to 3.

Losses in an AB exit nozzle are high – the flameholders and so forth act as obstructions. It is preferable to diffuse the flow speed at the AB from higher speed to Mach 0.2 to 0.3, which results in a small bulge in the jet-pipe diameter around that area. A combat aircraft fuselage must be able to house this bulge.

Typically, in this book, a long-duct nacelle is preferred to obtain better thrust and fuel economy and to offset the weight gain as compared to short-duct nacelles (see Figure 10.21). The pressure increase across the fan (i. e., secondary cold flow) is sub­stantially lower than the pressure increase of the primary airflow. The secondary airflow does not have the addition of heat as in the primary flow. The cooler and lower exit pressure of the fan exit – when mixed with the primary hot flow within the long duct – reduces the final pressure to lower than the critical pressure, favoring a perfectly expanded exit nozzle (pe = p»). Through mixing, there is a reduction in the jet velocity, which provides a vital benefit in noise reduction (see Chapter 15) to meet airworthiness requirements. The long-duct nacelle exit plane can be sized to expand perfectly.

Primary flow has a subscript designation of p and secondary flow has a sub­script designation of s. Therefore, Fp and Vep denote primary flow thrust and exit velocity, respectively, and Fs and Vsp denote bypass flow thrust and its exit velocity, respectively. Thrust (F) equations of perfectly expanded turbofans are computed separately for primary and secondary flows and then added to obtain the net thrust, F, of the engine (i. e., a perfectly expanded nozzle):

F = Fp + Fs = [(mp + mf) X Vep mp X V00] + [ms x (Ves 1»)]

Specific thrust in terms of primary flow becomes (f = fuel-to-air ratio), or:

F/mp = [(1 + f) x Vep + BPR x Ves – V» x (1 + BPR)] (10.12)

If the fuel flow is ignored, then:

F/mp = [Vep – V»] + BPR x (Ves – V») (10.13)

For kinetic energy (KE):

ke = mp^Vp – V2)] + ms[1h{VSp2 – V»2)] or (10.14)

KE/thp = [i/2(Vep2 – V2)] + BPR x [i/2(Vsp2 – V»2)]

At a given design point (i. e., flight speed V»), BPR, fuel consumption, and mp are held constant. Then, the best specific thrust and KE are found by varying the fan exit velocity for a given Vep, setting the differentiation relative to Ves equal to zero. (This may be considered as trend analysis for ideal turbofan engines; real engine analysis is more complex.)

Then, by differentiating Equation 10.13:

d( F/m p)/d(VeS) = 0 = d(Vep)/d(VeS)+BPR (10.15)

Equation 10.14 becomes:

d(KE/thp)/d(VeS) = 0 = Vepd(Vep)/d(VeS) + BPR x VSp (10.16)

Combining Equations 10.15 and 10.16:

-BPR x Vep + BPR x Vsp = 0

Because BPR = 0, the optimum is when:

Vep = Vsp (10.17)

That is, the best specific thrust is when the primary (i. e., hot core) exit-flow velocity equals the secondary (i. e., cold fan) exit-flow velocity.

Equation 10.4 gives the turbojet propulsive efficiency, np = v2VV, for a simple turbojet engine; however, for the turbofan, there are two exit-plane velocities – for the hot-core primary flow (Vep) and for the cold-fan secondary flow (Ves). There­fore, an equivalent mixed turbofan exit velocity (Veq) can substitute for Ve in the previous equation. Fuel-flow rates are minor and can be ignored. The equivalent turbofan exit velocity (Veq) is obtained by equating the total thrust (i. e., a perfectly expanded nozzle) as if it were a turbojet engine with total airmass flow (mp + ms).

Entropy, e

Figure 10.13. Afterburning turbojet and T-s diagram (real cycle)

Thus:

(mp + ms) X (Veq Vc») — mp X (Vep V(Xi’) + ms X (Ves Vc»)

or (1 + BPR) X (Veq – VTO) — (Vep – VTO) + BPR X (Ves – Vx)

or (1 + BPR) X Veq — (Vep – VTO) + BPR X (Ves – VTO) + VTO X (1 + BPR)

— Vep + BPR X Ves

or Veq — [Vep + BPR X Ves]/(1 + BPR) (10.18)

Then, turbofan propulsive efficiency:

2VTO

npf V V

eq + to

Large engines could benefit from weight savings by installing short-duct turbo­fans; some smaller aircraft also use short-duct nacelles.

In Figure 10.11, consider a CV (dashed line; note the waist-like shape of the simple turbojet) representing a straight-through, axi-symmetric turbojet engine. The CV and the component station numbers are as shown in the drawing and conventions in Figure 10.5; the gas turbine intake starts with the subscript 0, or to, and ends at the nozzle exit plane with subscript 5, or e. The free-stream airmass flow rate, ma, is inhaled into the CV at the front face perpendicular to the flow, the fuel-mass flow rate mf (from the onboard tank) is added at the combustion chamber, and the product flow rate (ma + m f) is exhaust from the nozzle plane perpendicular to flow. It is assumed that the inlet-face static pressure is pTO, which is fairly accurate. Precompression exists but, for the ideal consideration, it has no loss.

Flow does not cross the other two lateral boundaries of the CV because it is aligned with the walls of the engine. Force experienced by this CV is the thrust produced by the engine. Consider a cruise condition with an aircraft velocity of Vto. At cruise, the demand for air inhalation is considerably lower than at takeoff.

The intake area is sized between the two demands. At cruise, the intake-stream – tube cross-sectional area is smaller than the intake-face area – it is closer to that of the exit-plane area, A (the gas exits at a very high velocity). Because in an ideal condition there is no precompression loss, Station 0 may be considered to have free – stream properties with the subscript to.

From Newton’s second law, applied force F = rate of change of momentum + net pressure force (the momentum rate is given by the mass flow rate), where the inlet momentum rate = ma Vto and the exit momentum rate = (ma + m f )Ve.

Therefore:

the rate of change of momentum = (ma + mf )Ve – ma Vto (10.6)

The net pressure force between the intake and exit planes = peAe – pTO Ato (i. e., the axi-symmetric side pressure at the CV walls cancels out). Typically, at cruise, a sufficiently upstream Ae & Ato. Therefore:

F = (ma + mf )Ve – ma V» + Ae(pe – Pc») = net thrust (10.7)

Then:

(ma + mf )Ve + Ae(Pe – Pto) = gross thrust

and maVTO = ram drag (with – ve sign, it must be drag). It is the loss of energy seen as drag due to the slowing down of the incoming air as the ram effect. This gives:

net thrust = gross thrust – ram drag; Ae (pe – pTO) = pressure thrust

In general, subsonic commercial transport turbofans have a convergent nozzle, and the exit area is sized such that during cruise, pe & pTO (known as a perfectly expanded nozzle). This is different for military aircraft engines, especially with AB, when pe > pTO requires a convergent-divergent nozzle.

For a perfectly expanded nozzle, net thrust:

F = (ma + mf)Ve – ma V«, (10.8)

Further simplification is possible by ignoring the effect of fuel flow, m f, because m a » m f.

Then, the thrust for a perfectly expanded nozzle is:

At sea-level, static-takeoff thrust (TSLs) ratings V» = 0, which gives:

F = tha Ve+Ae(pe – p») (10.10)

Equation 10.10 indicates that the thrust increase can be achieved by increasing the intake airmass flow rate and/or increasing the exit velocity.

Equation 10.4 gives the propulsive efficiency:

Clearly, jet-propelled aircraft with low flight speeds have poor propulsive effi­ciency, np. Jet propulsion is favored for aircraft flight speeds above Mach 0.6.

The next question is: Where does the thrust act? Figure 10.12 shows a typical gas turbine engine in which the thrust is acting over the entire engine; the aircraft senses the net thrust transmitted through the engine-mounted bolts.

Figure 10.12 shows a typical straight-through turbojet pressure, velocity, and temperature variation along the length as airmass flows through. Readers may note the scale; within each component, the velocity change is negligible.

Figure 10.9a depicts a standard schematic diagram representing a simple straight – through turbojet engine, as shown in Figure 10.4, with appropriate station numbers. The thermodynamic cycle associated with gas turbines is known as the Joule cycle (also known as the Brayton cycle). Figure 10.9b is the corresponding temperature- entropy diagram of an ideal Joule cycle in which compression and expansion take place isentropically.

Real engine processes are not isentropic and losses are involved associated with increased entropy. Figure 10.10 is a comparison of real and ideal cycles.

10.1 Formulation and Theory: Isentropic Case

Gas turbine equations relevant to this book are provided in this section and are valid for all types of processes. For more details, see [2] through [6].

Most aircraft piston engines are the reciprocating type (i. e., positive displacement, intermittent combustion): The smaller ones have an air-cooled, two-stroke cycle; the

Figure 10.8. Aircraft piston engine with installation components

larger ones (typically, more than 200 HP) have a liquid-cooled, four-stroke cycle. There are a few rotary-type positive-displacement engines (e. g., Wankel) – attrac­tive in principle but they have sealing problems. Cost-wise, rotary-type positive- displacement engines are not yet popular; cost will decrease with increased produc­tion. Figure 10.8 shows an aircraft piston engine with installation components.

To improve high-altitude performance (with low air density), supercharging is used. Figure 10.8 shows a vane-supercharging type for precompression. Also, AVGAS differs slightly from MOGAS. Recently, some engines for the homebuilt category have been allowed to use MOGAS. Recently for small aircraft application, diesel fuel-powered piston engines have appeared in the market.

Piston engines are the oldest type used for powering aircraft. Over the life cycle of an aircraft, gas turbines are more cost effective for engine sizes of more than 500 HP. Currently, general-aviation aircraft are the main users of piston engines. Small recreational aircraft invariably are powered by piston engines.

Lower-speed aircraft can use propellers for thrust generation. Therefore, instead of driving a smaller encased fan (i. e., turbofan), a large propeller (i. e., turboprop) can be driven by a gas turbine engine to improve efficiency because the exhaust energy can be further extracted to a very low exhaust velocity (i. e., nearly zero noz­zle thrust). Some residual jet thrust is left at the nozzle exit plane when it needs to be added to the propeller thrust. The nozzle thrust is converted to HP and, together with the SHP generated, it becomes the equivalent SHP (ESHP).

However, a large propeller diameter limits rotational speed due to both aero­dynamic (i. e., transonic blade tips) and structural (i. e., centrifugal force) consider­ations. Heavy reduction gears are required to reduce the propeller rpm to a desir­able level. Propeller efficiency decreases when aircraft are operating at flight speeds above Mach 0.5. For shorter-range flights, a turboprop’s slower speed does not become time-critical to the users, yet it offers better fuel economy. Figure 10.7 is a schematic diagram of a typical turboprop engine. Modern turboprops have up to eight blades (see Figure 10.31), which allow a reduction of the diameter size and operate at a relatively higher rpm and aircraft speed.

Afterburning (AB) is another way of thrust augmentation intended exclusively for the supersonic combat aircraft category (the Concorde is the only civil aircraft that used AB). Figure 10.6 is a cutaway diagram of a modern AB engine intended for combat aircraft.

Figure 10.7. Schematic diagram of a turboprop engine

The simple straight-through turbojets have a relatively small frontal area result­ing in low drag and excess air in the exhaust flow. If additional fuel can be burned in the exhaust nozzle beyond the turbine exit plane, additional thrust can be gener­ated to propel an aircraft at a considerably higher speed and acceleration, thereby also possibly improving propulsive efficiency. However, the reason for using AB arises from the mission demand, such as at takeoff with a high payload and accelera­tion to engage or disengage during combat and evasion maneuvers. Mission demand overrides the fact that there is a high level of energy rejection in the high exhaust velocity. Fuel economy degrades with AB – it takes 80 to 120% more fuel burn to gain a 30 to 50% increase in thrust. Currently, most supersonic aircraft engines have some BPR when AB is done in the cooler mixed flow past the turbine section of the primary flow.

The energy extraction through the additional turbine lowers the rejected energy at the exhaust, resulting in a lower exhaust velocity, pressure, and temperature. The additional turbine drives a fan in front of the compressor. The large amount of air – mass flowing through the fan provides thrust. Part of the intake airmass flow through the fan is diverted (i. e., bypassed as the cold secondary flow) around the engine core and does not burn. The primary flow flows through the combustion chamber and is known also as the core flow or hot flow. Figure 10.5 is a schematic diagram of a turbofan engine (the top of the figure is a bare PW 4000). The lower exhaust veloc­ity reduces engine noise. At the design point (i. e., LRC), the lower exhaust pressure permits the nozzle exit area to be sized to make the exit pressure equal to the ambi­ent pressure (i. e., in a perfectly expanded nozzle). This is unlike simple turbojets, which can have a higher exit pressure.

Readers should note that the component-station-numbering system follows the same pattern as for the simple straight-through turbojet. The combustion chamber in the middle maintains the same numbers (i. e., 2-3). The only difference is the fan exit, which has the subscript f. The intermediate stages of the compressor and the turbine are primed.

Typically, the BPR (see Equation 10.2) for commercial jet-aircraft turbo­fans (i. e., high-subsonic flight speeds of less than Mach 0.98) is around 4 to 7. Recently, turbofans for the newer Boeing787, Airbus350 and Bombardier Cseries have reached BPR of 8 to 12. For military aircraft applications (i. e., supersonic flight speeds of up to Mach 2.5), the BPR is around 1 to 3. A lower BPR keeps the fan diameter smaller and, hence, lowers the frontal drag. Multispool drive shafts offer better efficiency and response characteristics, mostly with two concentric shafts. The shaft driving the low-pressure (LP) section runs inside the hollow shaft of the high – pressure (HP) section (see Figure 10.5). Three shaft turbofans have been designed, but most of the current designs use a twin spool. The recent advent of a geared turbofan is indicative of better fuel efficiencies.

nacelle

Shaded area is the HP module (compressor and turbine)

HP shaft goes through hollow LP shaft

Figure 10.5. Schematic diagram of a pod-mounted, long-duct, two-shaft turbofan engine

A lower fan diameter compared to the propeller permits higher rotational speed and provides the scope for a thinner aerofoil section to extract better aerody­namic benefits. The higher the BPR, the better is the fuel economy. A higher BPR demands a larger fan diameter when reduction gears may be required to keep the revolutions per minute (rpm) at a desired level. Ultra-high BPR (UHBPR) turbo­fans approach the class of a ducted-fan, ducted-propeller, or propfan engine. This type of engine has been built, but its cost versus performance has prevented it from breaking into the market.

This section describes various types of gas turbines and introduces piston engines. Aircraft propulsion depends on the extent of thrust produced by the engine. Sec­tion 10.11 presents the thrust and power available from various types of engines. Statistics for various types of aircraft engines are previous at the end of this chapter. Gas turbine sizes are progressing in making engines both larger and smaller than current sizes – that is, expanding the application envelope.

10.4.1 Simple Straight-Through Turbojet

The most elementary form of a gas turbine engine is a simple straight-through turbojet, shown schematically in Figure 10.4. In this case, the intake airflow goes straight through the entire length of the engine and exits at a higher velocity and temperature after the processes of compression, combustion, and expansion. This type of engine burns like a stove in a pressurized environment. Readers may note the “waisting” of the airflow passage as a result of the compressor reducing the volume as the turbine expands. Typically, at the LRC condition, the free-stream tube located far upstream is narrower in diameter than at the compressor face. As a result, airflow ahead of the intake plane slows down during the precompression phase.

Components associated with the thermodynamic processes within the engine have assigned station numbers, as listed in Table 10.3. (A bare engine does not have an intake and exhaust nozzle.)

Overall engine efficiency improves if the higher energy of the exhaust gas of a straight-through turbojet is extracted through an additional turbine, which can drive a fan in front of a compressor (i. e., for a turbofan engine) or a propeller (i. e., for a turboprop engine).