International Standards

The International Standard Atmosphere (ISA) defines standard day ambient temperature and pressure up to an altitude of 30,500 m (100,066 ft). The term ISA conditions alone would imply zero relative humidity.

US Military Standard 210 (MIL 210) is the most commonly used standard for defining likely extremes of ambient temperature versus altitude. This is primarily an aerospace standard, and is also widely used for land-based applications though with the hot and cold day temperature ranges extended. Table 19–1 shows the ambient pressure and temperature relationships of MIL 210 and ISA.

For land-based engines, performance data are frequently quoted at the single point ISO conditions, as stipulated by the International Organization for Standardization (ISO). These are:

Ambient Pressure and Pressure Altitude

Pressure altitude, or geo-potential altitude, at a point in the atmosphere is defined by the level of ambient pressure, as per the International Standard Atmosphere. Pressure altitude is therefore not set by the elevation of the point in question above sea level. For example, due to prevailing weather conditions a ship at sea may encounter a low ambient pressure of, say, 97.8 kPa, and hence its pressure altitude would be 300 m.

Table 19–1 includes the ISA definition of pressure altitude versus ambient pressure, and Figure 19–1 shows the relationship graphically. It will be observed that pressure falls exponentially from its sea level value of 101.325 kPa (14.696 psia) to 1.08 kPa (0.16 psia) at 30,500 m (100,066 ft).

The highest value of ambient pressure for which an engine would be designed is 108 kPa (15.7 psia). This would be due to local conditions and is commensurate with a pressure altitude of –600 m (–1968 ft).

Ambient Temperature

Table 19–1 also presents the ISA standard day ambient temperature, together with MIL 210 cold and hot day temperature, versus pressure altitude. Figure 19–2 shows these three lines of ambient temperature plotted versus pressure altitude.

Standard day temperature falls at the rate of approximately 6°C per 1000 m (2°C for 1000 ft) until a pressure altitude of 11,000 m (36,089 ft), after which it stays constant until 25,000 m (82,000 ft). This altitude of 11,000 m is referred to as the tropopause; the region below this is the troposphere, and that above it is the stratosphere. Above 25,000 m standard day temperature rises again.

The minimum MIL 210 cold day temperature of 185.9 K (–87.3°C) occurs between 15,545 m (51,000 ft) and 18,595 m (61,000 ft). The maximum MIL 210 hot day temperature is 312.6 K (39.5°C) at sea level.

Relative Density and the Speed of Sound

Relative density is the atmospheric density divided by that for an ISA standard day at sea level. Table 19–1 includes relative density, the square root of relative density, and the speed of sound for cold, hot and standard days. Figures 19–3 through 19–5 present this data graphically. These parameters are important in understanding the interrelationships between the different definitions of flight speed.

Density falls with pressure altitude such that at 30,500 m (100,066 ft) it is only 1.3% of its sISA sea level value. The maximum speed of sound of 689.0 kt (1276 km/h, 792.8 mph) occurs on a hot day at sea level. The minimum value is 531.6 kt (984.3 km/h, 611.6 mph), occurring between 15,545 m (51,000 ft) and 18,595 m (61,000 ft).

Specific and Relative Humidity

Atmospheric specific humidity is variously defined either as:

The former definition is used exclusively herein; for most practical purposes the difference is small anyway. Relative humidity is specific humidity divided by the saturated value for the prevailing ambient pressure and temperature.

Humidity has the least powerful effect upon engine performance of the three ambient parameters. Its effect is not negligible, however, in that it changes the inlet air’s molecular weight, and hence basic properties of specific heat and gas constant. In addition, condensation may occasionally have gross effects on temperature. Wherever possible humidity effects should be considered, particularly for hot days with high levels of relative humidity.

For most gas turbine performance purposes, specific humidity is negligible below 0°C, and also above 40°C. The latter is because the highest temperatures only occur in desert conditions, where water is scarce. MIL 210 gives 35°C as the highest ambient temperature at which to consider 100% relative humidity.

Figure 19–6 presents specific humidity for 100% relative humidity versus pressure altitude for cold, standard, and hot days. For MIL 210 cold days specific humidity is almost zero for all altitudes. The maximum specific humidity will never exceed 4.8%, which would occur on a MIL 210 hot day at sea level. In the troposphere (i.e., below 11,000 m) specific humidity for 100% relative humidity falls with pressure altitude, due to the falling ambient temperature. Above that, in the stratosphere, water vapor content is negligible, almost all having condensed out at the colder temperatures below.

Figures 19–7 and 19–8 facilitate conversion of specific and relative humidities. Figure 19–7 presents specific humidity versus ambient temperature and relative humidity at sea level. For other altitudes Figure 19–8 presents factors to be applied to the specific humidity obtained from Figure 19–7. For a given relative humidity specific humidity is higher at altitude because whereas water vapor pressure is dependent only on temperature, air pressure is significantly lower.

Industrial Gas Turbines

The environmental envelope for industrial gas turbines, both for power generation and mechanical drive applications, is normally taken from Table 19–1 up to a pressure altitude of around 4500 m (or 15,000 ft). Hot and cold day ambient temperatures beyond those of MIL 210 are often used, ±50°C being typical at sea level. For specific fixed locations, altitude is known, and S curves are available defining the annual distribution of ambient temperature: these allow lifing assessments and rating selection. (The name derives from the characteristic shape of the curve, which plots the percentage of time for which a particular temperature level would be exceeded.)

The range of specific humidities for an industrial gas turbine would be commensurate with 0–100% relative humidity over most of the ambient temperature range, with some alleviation at the hot and cold extremes.

Automotive Gas Turbines

Most comments are as per industrial engines, except that narrowing down the range of ambient conditions for a specific application based on a fixed location is not appropriate.

Marine Gas Turbines

The range of pressure altitudes at sea is governed by weather conditions only, as the element of elevation that significantly affects all other gas turbine types is absent. The practice of the US Navy, the most prolific user of marine gas turbines, is to take the likely range of ambient pressure as 87–108 kPa (12.6–15.7 psia). This corresponds to a pressure altitude variation of –600–1800 m (–1968–5905 ft).

At sea, free stream air temperature (i.e., that not affected by solar heating of the ship’s decks) matches sea surface temperature, day or night. Owing to the vast thermal inertia of the sea there is a significant reduction in the range of ambient temperature that marine gas turbines encounter when on the open sea relative to land based or aircraft gas turbines. However, ships must also be able to operate close to land, including polar ice fields and the Persian Gulf. Consequently for operability (if not lifing) purposes, a wide range of ambient temperature would normally be considered for a marine engine. The most commonly used ambient temperature range is that of the US Navy, which is –40–50°C. US Navy ratings are proven at 38°C, giving some margin on engine life.

The relative humidity range encountered by a marine gas turbine would be unlikely to include zero, due to the proximity of water. In practice values above 80% are typical. The upper limit would be commensurate with 100% relative humidity over most of the ambient temperature range, again with some alleviation at the hot and cold extremes.

Aircraft Engines

The environmental envelope for aircraft engines is normally taken from Table 19–1 up to the altitude ceiling for the aircraft. The specific humidity range is that corresponding to zero to 100% relative humidity as per Figure 19–6.

Industrial Engines

Overall installation inlet pressure loss due to physical ducting, filters and silencers is typically 100 mm H₂O at high power. Installation exhaust loss is typically 100–300 mm H₂O (0.981 kPa, 0.142 psi to 2.942 kPa, 0.427 psi); the higher values occur where there is a steam plant downstream of the gas turbine.

Automotive Engines

In this instance both installation inlet and exhaust loss are typically 100 mm H₂O (0.981 kPa, 0.142 psi).

Marine Engines

Installation intake and exhaust loss values at rated power may be up to 300 mm H₂O (2.942 kPa, 0.427 psi) and 500 mm H₂O (4.904 kPa, 0.711 psi), respectively, dependent upon ship design. Standard values used by the US Navy are 100 mm H₂O (0.981 kPa, 0.142 psi) and 150 mm H₂O (1.471 kPa, 0.213 psi).

Aircraft Engines

For a pod mounted turbofan cruising at 0.8 Mach number, the total pressure loss from free stream to the flight intake/engine intake interface due to incomplete ram recovery and the installation intake may be as low as 0.5% ΔP/P, whereas for a ramjet operating at Mach 3 the loss may be nearer 15%. For a helicopter engine buried behind filters the installation intake total pressure loss may be up to 2%, and there may also be an installation exhaust pressure loss due to exhaust signature suppression devices.

Typical Flight Envelopes for Major Aircraft Types

Aircraft engines must operate at a range of forward speeds in addition to the environmental envelope. The range of flight Mach numbers for a given altitude is defined by the flight envelope. Figure 19–9 presents typical flight envelopes for the seven major types of aircraft.

For each flight envelope the minimum and maximum free stream temperatures and pressures the engine would experience are shown, together with basic reasons for the shape of the envelope. Where auxiliary power units are employed the same free stream conditions are experienced as for the propulsion unit. The intake ram recovery is often lower, however, due both to placement at the rear of the fuselage and drag constraints on the intake design.

Free Stream Total Pressure and Temperature

The free stream total pressure (P0) is a function of both pressure altitude and flight Mach number. Free stream total temperature (T0) is a function also of ambient temperature and flight Mach number. Both inlet pressure and temperature are fundamental to engine performance. They are often used to refer engine parameters to ISA sea level static conditions, via quasi dimensionless parameter groups. To do this the following ratios are defined:

$\text{DELTA}\left(\text{δ}\right)=\text{P0/}101\text{.}325\text{kPa}$

$\text{THETA}\left(\text{θ}\right)=\text{T0/}288\text{.}15\text{K}$

Definitions of Flight Speed

Traditionally aircraft speed has been measured using a pitot-static head located on a long tube projecting forward from the wing or fuselage nose. The difference between total and static pressure is used to evaluate velocity, which is shown on a visual display unit or gauge in the cockpit. The device is normally calibrated at sea level, which has given rise to a number of definitions of flight speed:

VEAS and VCAS are functions of pressure altitude and Mach number only. The difference between VEAS and VCAS is termed the scale altitude effect (SAE), and is independent of ambient temperature. Conversely, VTAS is a function of ambient temperature and Mach number, which is independent of pressure altitude.

To a pilot both Mach number and ground air speed are important. The former dictates critical aircraft aerodynamic conditions such as shock or stall, whereas the latter is vital for navigation. For gas turbine engineers Mach number is of paramount importance in determining inlet total conditions from ambient static. Often, however, when analyzing engine performance data from flight tests only VCAS or VEAS are available.

Equation of State for a Perfect Gas

A perfect gas adheres to Formula F19.1. All gases employed as the working fluid in gas turbine engines, except for water vapor, may be considered as perfect gases without compromising calculation accuracy. When the mass fraction of water vapor is less than 10%, which is usually the case when it results from the combination of ambient humidity and products of combustion, then for performance calculations the gas mixture may still be considered perfect. When water vapor content exceeds 10% the assumption of a perfect gas is no longer valid and for rigorous calculations steam tables must be employed in parallel, for that fraction of the mixture.

A physical description of a perfect gas is that its enthalpy is only a function of temperature and not pressure, as there are no intermolecular forces to absorb or release energy when density changes.

Molecular Weight and the Mole

The molecular weight for a pure gas is defined in the Periodic Table. For mixtures of gases, such as air, the molecular weight may be found by averaging the constituents on a molar (volumetric) basis. This is because a mole contains a fixed number of molecules, as described.

A mole is the quantity of a substance such that the mass is equal to the molecular weight in grams. For any perfect gas 1 mole occupies a volume of 22.4 liters at 0°C, 101.325 kPa. A mole contains the Avogadro’s number of molecules, 6.023 × 10²³.

Specific Heat at Constant Pressure (CP) and at Constant Volume (CV)

These are the amounts of energy required to raise the temperature of 1 kilogram of the gas by 1°C, at constant pressure and volume, respectively. For gas turbine engines, with a steady flow of gas (as opposed to piston engines where it is intermittent) only the specific heat at constant pressure, CP, is used directly. This is referred to hereafter simply as specific heat.

For the gases of interest specific heat is a function of only gas composition and static temperature. For performance calculations total temperature can normally be used up to Mach numbers of 0.4 with negligible loss in accuracy, since dynamic temperature remains a low proportion of the total.

Gas Constant (R)

The gas constant appears extensively in formulae relating pressure and temperature changes, and is numerically equal to the difference between CP and CV. The gas constant for an individual gas is the universal gas constant divided by the molecular weight, and has units of J/kg K. The universal gas constant has a value of 8314.3 J/mol K.

Ratio of Specific Heats, Gamma (γ) (Formulae F19.6–F19.8)

This is the ratio of the specific heat at constant pressure to that at constant volume. Again, it is a function of gas composition and static temperature, but total temperature may be used when the Mach number is less than 0.4. Gamma appears extensively in the “perfect gas” formulae relating pressure and temperature changes and component efficiencies.

Dynamic Viscosity (VIS) and Reynolds Number (RE) (Formulae F19.9)

Dynamic viscosity is used to calculate the Reynolds number, which reflects the ratio of momentum to viscous forces present in a fluid. The Reynolds number is used in many performance calculations, such as for disc windage, and has a second-order effect on component efficiencies. Dynamic viscosity is a measure of the viscous forces and is a function of gas composition and static temperature. As viscosity has only a second-order effect on an engine cycle, total temperature may be used up to a Mach number of 0.6. The effect of fuel air ratio (gas composition) is negligible for practical purposes.

The units of viscosity of N s/m² are derived from N/(m/s)/m; force per unit gradient of velocity. Gas velocity varies in a direction perpendicular to the flow in the boundary layers on all gas washed surfaces.

Total or Stagnation Temperature (T) (Formula F19.10)

Total temperature is the temperature resulting from bringing a gas stream to rest with no work or heat transfer. Note that here “at rest” means relative to the engine, which may have a flight velocity relative to the Earth. The difference between the total and static temperatures at a given point is called the dynamic temperature. The ratio of total to static temperature is a function of only gamma and Mach number, as per Formula F19.10.

In general for gas turbine performance calculations total temperature is used through the engine, evaluated at engine entry from the ambient static temperature and any ram effect. At locations between engine components total temperature is a valid measure of energy changes. In addition, this aids comparison between predictions and test data, as it is only practical to measure total temperature. For most component design purposes, however, static conditions are also relevant, as for example the Mach number is often high (1.0 and greater) at entry to a compressor stator or turbine rotor blade.

Total temperature is constant for flow along ducts where there is no work or heat transfer, such as intake and exhaust systems. Total and static temperature diverge much less rapidly versus Mach number than do total and static pressure, as described below.

Total or Stagnation Pressure (P) (Formulae F19.11, F19.12, and F19.13)

Total pressure is that which would result from bringing a gas stream to rest without any work or heat transfer, and without any change in entropy. Total pressure is therefore an idealized property.

The difference between total and static pressure at a point is called either the dynamic pressure, dynamic head, or velocity head (Formulae F19.12 and F19.13). The term head relates back to hydraulic engineering. The ratio of total to static pressure, as for temperature, is a function of only gamma and Mach number. Most performance calculations are conducted using total pressure, that at engine inlet again resulting from ambient static plus intake ram recovery.

Total pressure is not constant for flow through ducts, being reduced by wall friction and changes in flow direction, which produce turbulent losses. Both these effects act on the dynamic head; the pressure loss in a duct of given geometry and inlet swirl angle is almost always a fixed number of inlet dynamic heads. For this reason for performance calculations both the total and static pressure must often be evaluated at entry to ducts. Again, for component design purposes both the total and static values are of interest.

Total and static pressure diverge much more rapidly versus Mach number than do total and static temperature. Calculation of pressure ratio from temperature ratio is far more sensitive to errors in the assumption of the mean gamma than the reverse calculation.

Specific Enthalpy (H) (Formulae F19.14, F19.15)

This is the energy per kilogram of gas relative to a stipulated zero datum. Changes in enthalpy, rather than absolute values, are important for gas turbine performance. Total or static enthalpy may be calculated, depending on which of the respective temperatures is used. Total enthalpy, like total temperature, is most common in performance calculations.

Specific Entropy (S)

Traditionally the property entropy has been shrouded in mystery, primarily due to being less tangible than the other properties discussed in this section. Entropy relates to other thermodynamic properties relevant to gas turbine performance, and thereby helps overcome these difficulties.

During compression or expansion the increase in entropy is a measure of the thermal energy lost to friction, which becomes unavailable as useful work.

Dry Air

Dry air comprises the elements in Table 19–2.

There are also trace amounts of helium, methane, krypton, hydrogen, nitrous oxide, and xenon. These are negligible for gas turbine performance purposes.

Combustion Products

When a hydrocarbon fuel is burned in air, combustion products change the composition significantly. Atmospheric oxygen is consumed to oxidize the hydrogen and carbon, creating water and carbon dioxide, respectively. The degree of change in air composition depends both on fuel air ratio and fuel chemistry. The fuel air ratio such that all the oxygen is consumed is termed stoichiometric.

Distilled liquid fuels such as kerosene or diesel each have relatively fixed chemistry. Properties of their combustion products can be evaluated versus fuel air ratio and temperature using unique formulae, with the fuel chemistry inbuilt. In contrast, the chemistry of natural gas varies considerably. All natural gases have a high proportion of light hydrocarbons, often with other gases such as nitrogen, carbon dioxide, or hydrogen. Because the composition of natural gas combustion products varies, along with the fuel chemistry, unique formulae for their gas properties do not exist, hence the calculation is more complex.

Constant, Standard Values for CP and Gamma

This normally uses the following approximations:

This is the least accurate method, giving errors of up to 5% in leading performance parameters. It should only be used in illustrative calculations for teaching purposes, or for crude “ballpark” estimates.

Values for CP and Gamma Based on Mean Temperature

For formulae using CP and gamma it is most accurate to base these values on the mean temperature within each component, i.e., the arithmetic mean of the inlet and exit values. It is less accurate to evaluate CP and gamma at inlet and exit, and then take a mean value for each.

For dry air and combustion products of kerosene or diesel the formulae given for CP as a function of temperature and fuel air ratio give accuracies of within 1.5% for leading performance parameters. The largest errors occur at the highest pressure ratios.

Specific Enthalpy and Entropy—Dry Air, and Diesel or Kerosene

For fully rigorous calculations changes in enthalpy and entropy across components must be accurately evaluated. This improves accuracy to be within 0.25% for leading parameters at all pressure ratios. Here polynomials of specific enthalpy and entropy are utilized, obtained by integration of the standard polynomials for specific heat. In these methods, a formula for specific heat is therefore still required.

The use of specific enthalpy and entropy for performance calculations is now almost mandatory for computer “library” routines in large companies.

Specific Enthalpy and Entropy—Natural Gas

For the combustion products of natural gas it is logical to use specific enthalpy and entropy only if CP is evaluated accurately.

Molecular Weight and Gas Constant

Data for gases of interest are listed in Table 19–3.

Figure 19–10 shows the gas constant resulting from the combustion of leading fuel types in air plotted versus fuel air ratio. It is not possible to provide all encompassing data for natural gas due to the wide variety of blends that occur. For indicative purposes combustion of a sample natural gas has been used. The following are apparent:

The effect of gas fuel is more powerful than that of liquid fuels, primarily due to the constituent hydrocarbons being lighter (i.e., containing less carbon and more hydrogen); this results in a higher proportion of water vapor after combustion, which has a significantly lower molecular weight than the other constituents.

Specific Heat and Gamma

Figures 19–11 and 19–12 present specific heat and gamma respectively for dry air and combustion products versus static temperature and fuel air ratio for kerosene or diesel fuels.

Figure 19–13 shows the ratio of specific heat following the combustion of the sample natural gas to that for kerosene, versus fuel to air ratio. This plot is sensibly independent of temperature. As stated earlier, specific heat is noticeably higher following the combustion of natural gas due to the higher resultant water content, which significantly impacts engine performance.

Figure 19–14 shows specific heat respectively versus temperature for the individual gases present in air and combustion products. The higher value for water vapor is immediately apparent. For inert gases such as helium, argon, and neon, specific heat and gamma do not change with temperature.

Temperature Entropy Diagram for Dry Air

Most heat engine cycles are taught at university level via schematic illustration on a temperature–entropy (T–S) diagram. This approach becomes laborious to extend to “real” engine effects such as internal bleeds and cooling flows, but remains a useful indication of the overall thermodynamics of a known engine cycle. Figure 19–15 presents an actual temperature–entropy diagram for dry air, complete with numbers, showing lines of constant pressure. Such a diagram is rare in the open literature.

The following are important:

Entropy may be defined as thermal energy not available for doing work. In real compressors and turbines some energy goes into raising entropy, as some pressure is lost to real effects such as friction. The ideal work would be required or produced if entropy did not change, i.e., the process were isentropic. Isentropic efficiency is defined as the appropriate ratio of actual and ideal work, and is always less than 100%. (The term adiabatic efficiency is also commonly used, but is strictly incorrect. It only excludes heat transfer but not friction, and an isentropic process would have neither.)

Gas turbine cycles utilize the above processes, and rely on one other vital, fundamental thermodynamic effect:

Work input, approximately proportional to temperature rise, for a given compression ratio from low temperature is significantly lower than the work output from the same expansion ratio from higher temperature.

This is because on the T–S diagram lines of constant pressure diverge with increasing temperature and entropy.

This can be seen by considering a sample compression, heating, and expansion between two lines of constant pressure, using Figure 19–15. At an entropy value of 1.5 kJ/kg K the temperature rise required to go from 100 to 5100 kPa is 500 K. If fuel is now burned at this pressure level such that entropy increases to 2.75 kJ/kg K, and temperature to 1850 K, an expansion back to 100 kPa will achieve a temperature drop of around 1000 K. This clearly illustrates the rationale behind the Brayton cycle.

Schematic T–S Diagrams for Major Engine Cycles

Figures 19–16 through 19–21 show the key cycles of interest to gas turbine engineers.

Figure 19–16 shows the Carnot cycle. This is the most efficient cycle theoretically possible between two temperature levels. Gas turbine engines necessarily do not use the Carnot cycle, as unlike steam cycles they cannot add or reject heat at constant temperature.

Figure 19–17 shows the Brayton cycle. This is the basic cycle utilized by all gas turbine engines where heat is input at constant pressure. The effect of component inefficiency is shown by the non-vertical compression and expansion lines, a further difference from the ideal Carnot cycle. The form of the Brayton cycle is modified for heat exchangers and bypass flows.

Figure 19–18 presents the cycle for a turbofan. The bypass stream only undergoes partial compression, and no heating before expansion back to ambient pressure.

Figure 19–19 shows a heat exchanged cycle. Waste heat from exhaust gases is used to heat air from compressor delivery prior to combustion, thereby reducing the required fuel flow.

Figure 19–20 presents an intercooled cycle, where heat is extracted downstream of an initial compressor. This reduces the work required to drive a second compressor, and thereby increases power output.

Figure 19–21 shows a Rankine cycle with superheat. This is used in combined-cycle applications, with the gas turbine exhaust gases providing heat to raise steam. Where heat is added at constant temperature during evaporation a close approximation to the Carnot cycle is achieved, the main deviation being the non-ideal component efficiencies.

The Engine Design Point and Design Point Performance

For initial definition work, the operating condition where an engine will spend most time has been traditionally chosen as the engine design point. For an industrial unit this would normally be ISO base load, or for an aeroengine cruise at altitude on an ISA day. Alternatively some important high power condition may be chosen. Either way, at the design point the engine configuration, component design, and cycle parameters are optimized. The method used is the design point performance calculation. Each time input parameters are changed and this calculation procedure is repeated, the resulting change to the engine design requires a different engine geometry, at the fixed operating condition.

For the concept design phase described here the component design points are usually at the same operating condition as the engine design point. In a detailed design phase, however, this may not be true. For example, in the detailed design phase an aeroengine fan may be designed at the top of climb, the highest referred speed and flow, whereas the engine design point would be cruise. The term design point refers to the engine design point in the concept design phase, which is taken to be coincident with the component design points.

Engine Performance Parameters

A number of key parameters that define overall engine performance are utilized to assess the suitability of a given engine design to the application, or compare several possible engine designs. These engine performance parameters are described as follows:

For aircraft thrust engines there are a number of secondary performance parameters. These do not in themselves describe overall engine performance, but do help the engine designer understand the variation of the primary performance parameters across a number of designs. These parameters are as follows:

Thrust SFC is directly proportional to flight speed, and inversely proportional to both thermal and propulsive efficiencies. The dependency on the efficiencies is intuitively obvious. The choice of pressure ratio and SOT for minimum SFC is a compromise between maximizing propulsive and thermal efficiency.

The dependency of thrust SFC on flight speed arises because fuel flow relates to power with the thermal and propulsive efficiencies fixed, and propulsive power is directly proportional to flight speed for a given thrust.

Cycle Design Parameters

The fundamental thermodynamics of gas turbine cycles in relation to temperature entropy diagrams show that the changes in pressure and temperature that the working fluid experiences strongly affect the engine performance parameters. The degree of change of pressure and temperature are reflected via the following cycle design parameters.

Overall Pressure Ratio

This is total pressure at compressor delivery divided by that at the engine inlet.

Stator Outlet Temperature (SOT)

This is the temperature of the gas able to do work at entry to the first turbine rotor. Other terms are also used to reflect maximum temperatures in a cycle:

The standard definition for SOT, used herein, is:

Hence, in this definition, nozzle guide vane or platform cooling air entering upstream of the throat, or trailing edge cooling air ejected with the same momentum and direction as the main flow, would be included. However, front disc face cooling airflow would not be considered in evaluating SOT since it will not do work in the turbine rotor. For most gas turbine engine types it is desirable to raise SOT to as high a level as possible within mechanical design constraints.

In addition for turbofan engines two further cycle design parameters occur due to the parallel gas paths.

Fan Pressure Ratio

This is the ratio of fan delivery total pressure to that at fan inlet, and is usually lower for the core stream than the bypass due to lower blade speed.

Bypass Ratio

This is the ratio of mass flow rate for the cold stream to that for the hot stream.

Component Performance Parameters

A plethora of parameters define component performance in terms of efficiency, flow capacity, pressure loss, etc. As the level of component performance parameters improves, at fixed values of cycle design parameters, then the design point engine performance parameters also improve.

Changes in component performance parameters have a secondary effect on the optimum values of engine cycle design parameters.

Mechanical Design Parameters

For a given performance design point to be practical the mechanical design parameters must be kept within the limits of the materials, manufacturing, and production technology available. These are mechanical design restraints in relation to gas turbine engines:

Life Parameters

The two major life parameters are:

The TBO is governed mainly by creep and oxidation life, while cyclic life is dictated by thermal stress levels. Typical life requirements for the major gas turbine applications are as shown in Table 19–4.

Fuel Type

Kerosene is the standard aviation fuel while marine engines burn diesel and most industrial applications use natural gas. The highly distilled forms of diesel used make little difference to performance compared with kerosene, but natural gas gives performance improvements because of the higher resulting specific heat of the combustion products.

Design Point Diagrams

A design point diagram is created by plotting engine performance parameters versus the cycle parameters. To produce the diagram the design point calculations are repeated varying each cycle parameter in turn through the range of interest; every point is a different engine geometry. These diagrams are useful in the engine design process for selecting the optimum cycle parameters, or for comparing the performance of different gas turbine types. Initially design point diagrams are prepared using constant component performance parameter levels for all combinations of cycle parameters. Later in the engine design process different individual component performance levels are used for each combination of cycle parameters, based on aerothermal component design work.

Referred Parameters

The design point charts may be applied to any altitude if the referred form of specific power or thrust, SFC, etc. is used. In this way the datum values for the charts are adjusted. A change in flight Mach number does change engine matching, however, unless all nozzles are choked, and also changes ram drag. The charts are not directly adaptable to other flight Mach numbers.

Simple Cycle

This is the basic shaft power cycle. First, air is compressed, and then it is heated by burning fuel in the combustor. Next, expansion through one or more turbines produces power in excess of that required to drive the compressors, which is available as output. Whether the engine is single spool or free power turbine is only reflected in a design point diagram by any small change in assessed overall turbine efficiency:

Recuperated Cycle

This is as per the simple cycle except that a heat exchanger transfers some of the heat in the exhaust to the compressor delivery air. If the gas and air streams do not mix the heat exchanger is known as a recuperator; if they do mix, as with the automotive ceramic rotating matrix variety, it is known as a regenerator:

Intercooled Cycle

In this configuration the temperature of the air is reduced part way through the compression process using a heat exchanger and an external medium such as water. This increases power output via reduced compressor work, which is directly proportional to compressor inlet temperature. Generally thermal efficiency reduces as more fuel is required to reach a given SOT:

Intercooled and Recuperated Cycle

This is the combination of intercooling and recuperation. The thermal efficiency loss due to intercooling is offset by increased exhaust heat recovery due to the lower recuperator air-side entry temperature:

Combined Cycle

This is where a steam (Rankine) cycle also generates power, using exhaust heat from a simple cycle gas turbine. In addition, supplementary firing may be employed whereby a separate boiler supplements the gas turbine heat output when necessary. The curves are for a triple pressure reheated steam plant:

Turbojets

Here a simple cycle configuration produces thrust via high exhaust gas velocities.

Turbofans

Here air from the first compressor of what is otherwise a turbojet engine bypasses the remaining compressor and turbine stages. This air produces thrust directly, additional to that from the hot exhaust gases.

Ramjets

Here fuel is burned in air compressed solely by the intake ram effect; there is no turbomachinery. Engine performance parameters are therefore a function only of SOT and flight Mach number.

Statement of Requirements

A statement of requirements or specification for a new engine may be presented by a customer organization, or may be specified within the cycle designer’s company to meet a perceived market need. Realism is required in terms of technical challenge as well as potential development and unit production costs. The document must not just reflect what “Marketing” or “the Customer” desires but also what “Engineering” can achieve. It should be written jointly.

With respect to performance the statement of requirements must contain a minimum of:

“First Cut” Design Points

Generic design point diagrams will show what engine configurations, together with the levels of cycle parameters, are likely to meet the statement of requirements. Next, initial component performance targets are set. For the narrow range of cycle parameters of interest, refined design point diagrams are then produced. Usually component efficiency variation with engine cycle parameters is unknown at this stage. As far as possible the cycle designer takes account of mechanical design limits such as temperature levels and considers the use of existing or scaled components.

From these design point diagrams a first cut design point for each engine configuration under consideration is formally issued to component designers. An example of multiple configurations is that both simple cycle and recuperated engines may be competitive for a shaft power application.

For clarity all design points should have an identifying version number, as many more updates will follow in what is a highly iterative design process.

“First Cut” Aero-Thermal Component Design

The component specialists now reassess the component performance inputs assumed in the design points, and perform initial component sizing. Their analysis uses component design computer codes and empirical data. Mechanical design parameters are monitored and where there is doubt regarding limits mechanical technologists must be consulted.

Proper choice of rotational speed is crucial. It is key to achieving the target compressor and turbine efficiency levels as well as number of stages and diameter, and the compressor and turbine designers must agree upon it. The practicality or otherwise of using the number of spools that the cycle designer had in mind when setting the initial design points will soon be apparent, and weight and cost implications will become clear.

Component performance levels are now bid back to the performance engineer for each design point.

Design Point Calculations and Aero-Thermal Component Design Iterations

By repeating the design point calculations using the new bid component performance levels the concept design team can decide which engine configuration is the most suitable. If a bid is changed even for only one component, it will change the design point requirements for all others.

The process is repeated a number of times for the chosen configuration so that a number of cycle parameter combinations, each with specific component performance levels, can be compared.

Engine Layout

It is important to commence engine layout drawings as early as possible to highlight potential difficulties. For example, for a turbofan there would normally be a rigid engine diameter constraint. The core diameter required by the compressor and turbine designs, combined with the bypass duct annulus area to keep Mach number to a level where pressure loss is acceptable, may exceed this limit. In this instance the cycle designer may have to reduce bypass ratio, or the component designers increase rotational speed.

Off-Design Performance

Once a good candidate design point is emerging from the above iterations the cycle designer must then freeze the engine geometry and issue off-design performance. This should cover other key operating points as well as all corners of the operational envelope. Often at this juncture the component designers will not have created off-design maps for the “frozen” component designs so maps from existing components of similar design must be used, with factors and deltas applied to align them to the design point. The performance engineer must decide how the engine will be controlled for key ratings before running the off-design cases.

It is good practice at this stage for the performance engineer to highlight the operational cases where the maximum and minimum value of all key parameters occur, and any margins that need to be applied for say operating temperatures for a minimum engine.

Performance, Aero-Thermal, and Mechanical Design Investigation

The off-design performance data must be examined in detail by all parties. The cycle designer must check whether the engine design meets other performance requirements besides that at the design point. The aero-thermal and mechanical component designers must ensure that the components are satisfactory throughout the operational envelope with respect to component performance levels, stress, and vibration.

Iterations

A number of iterations through the whole process concludes the concept design phase. The resulting engine design is then compared with the statement of requirements and a judgment made as to whether to proceed into a detailed design and development program.

Minimum Engine

Owing to manufacture and build tolerances there will be significant engine-to-engine variation in production. Usually the engine designers deal with average engine data, that is, the average performance of all production engines. However, if the customer has been guaranteed a given performance level then even the minimum engine must achieve this standard. Hence a design margin must be built into the average engine performance targets. The shortfall in power or thrust and SFC of a minimum engine in a production run will be 1–3% versus the average engine if run at constant temperature, depending on build quality and engine complexity. If the customer has been guaranteed levels of power or thrust then this minimum engine will run up to 20 K hotter than an average engine to achieve power or thrust. This must be allowed for in the mechanical design.

Growth Potential

Almost any engine must be capable of adaptation to higher power or thrust levels. In aircraft applications the airframe usually requires it, intentionally or otherwise. In addition, addressing a wider market with development costs spread over a family of engines is attractive economically. Examples for turbofans include the Rolls Royce RB211/Trent family (thrust growth of 180 to 400+ kN in 25 years), the Rolls Royce civil and military Spey family, and the General Electric CF6 range. Margin is required on levels of temperature and speed, and there must be a practical route to increasing engine mass flow.

Engine Deterioration

At the concept design stage it is unusual to specify performance requirements after deterioration over a number of hours in service. However, if this is the case then a margin must be allowed for from the outset. The amount required depends on whether in service the engine is governed to fixed levels of temperature, speed or thrust/power. Depending on engine complexity typical margins at 10,000 hours, expressed at constant SOT, are: +3 to +6% SFC; –5 to –15% power.

Performance degradation is usually exponential with time, and after 10,000 hours there is minimal further deterioration. Practices vary regarding the accounting or not of running in effects caused by initial seal cutting as part of deterioration. The aim of running in is to improve performance levels by ensuring a gentle first contact on key seals, which removes less material than would a first in-service rub.

One further effect is compressor fouling where accretion of airborne particles on aerofoil surfaces reduces flow and efficiency, raising engine temperature levels. This applies mainly to ground based engines and is recoverable via compressor washing, for which appropriate fluid injection nozzles must be provided. The temperature increase depends entirely on site location, filter effectiveness, and washing frequency.

Installation Losses

Usually performance targets are stipulated as uninstalled, with the engine performance quoted from the engine intake flange to the engine exhaust or propelling nozzle flange. If the performance targets are installed then the magnitude of all installation effects must be stated, as follows:

Two case histories and one Mach number-altitude chart follow. Although the cases involve aircraft engine design, the principles are applicable to all gas turbines. The cases were written in 1972 and 1988, respectively, but they still draw interest today. Because the 1972 case is now hard to access, it is reprinted almost entirely.

Constant Static Pressure Mixing

If the upper boundary of the flow shown in Figure 19–23 is assumed to be flexible, then coolant and mainstream flows may be mixed under the assumption of a constant static pressure. Applying the equations of continuity, streamwise momentum, energy, and assuming thermally and calorically perfect gases for both mainstream and coolant, the result

$\frac{{\text{ΔP}}_{\text{t}}}{{\text{P}}_{{\text{t}}_{\text{g}}}}={\left(\frac{1+\frac{\text{γ}-1}{2}{\text{M}}_{\text{f}}^2}{1+\frac{\text{γ}-1}{2}{\text{M}}_{\text{g}}^2}\right)}^{\frac{\text{γ}-1}{\text{γ}}}-1$ (19.3)

(19.3)

${\text{M}}_{\text{f}}^2={[\frac{{\text{c}}_{\text{P}}(\text{γ}-1)(1+{\text{ε})(\text{T}}_{{\text{t}}_{\text{g}}}+{\text{εT}}_{{\text{t}}_{\text{c}}})}{{{(\text{V}}_{\text{g}}+{\text{εV}}_{\text{c}}\text{cos}\mathrm{\varphi })}^2}-\frac{(\text{γ}-1)}{2}]}^{-1}$ (19.4)

(19.4)

is obtained.

A Linear Approximation

The one-dimensional mixing of Figure 19–23 may also be approximated by adopting the “influence coefficient” method of Shapiro in the following manner.

For the compressible flow of a perfect gas having constant external and frictional forces, the differential variation of total pressure may be expressed as

${\text{dP}}_{\text{t}}=\left(\frac{\partial {\text{P}}_{\text{t}}}{\partial \text{A}}\right)\text{dA}+\left(\frac{\partial {\text{P}}_t}{\partial {\text{T}}_t}\right){\text{dT}}_{\text{t}}+\left(\frac{\partial {\text{P}}_t}{\partial \text{W}}\right)\text{dW}$

with partial derivatives of

$\begin{array}{l}\frac{\partial {\text{P}}_{\text{t}}}{\partial \text{A}}=0\\ \frac{\partial {\text{P}}_{\text{t}}}{\partial {\text{T}}_{\text{t}}}=-\frac{{\text{γM}}^2{\text{P}}_{\text{t}}}{{2\text{T}}_{\text{t}}}\\ \frac{\partial {\text{P}}_t}{\partial \text{W}}=-\frac{{\text{γM}}^2{\text{P}}_{\text{t}}}{\text{W}}\left(1-\frac{{\text{V}}_{\text{c}}}{\text{V}}\cos {\mathrm{\varphi }}_{\text{c}}\right)\frac{\text{dW}}{\text{W}}\end{array}$

as obtained from the derivation by Shapiro. Rearranging gives

$\frac{{\text{dP}}_{\text{t}}}{{\text{P}}_{\text{t}}}=\frac{-\text{γ}}{2}{\text{M}}^2\frac{{\text{dT}}_{\text{t}}}{{\text{T}}_{\text{t}}}-{\text{γM}}^2\left(1-\frac{{\text{V}}_{\text{c}}}{{\text{V}}_{\text{g}}}\cos {\mathrm{\varphi }}_{\text{c}}\right)\frac{\text{dW}}{\text{W}}$ (19.5)

(19.5)

which assumes that the injected fluid enters the mainstream at the local static pressure. Exact solution of Equation 19.5 requires integration of the functions, but, assuming that relatively small variations of T_t and W occur, the differentials may be considered Δ’s. The term “linear approximation” is thus derived from the fact that the equation is thus linear in the independent variables dT_t/T_t and dW/W. The term dT_t/T_t may be simplified by assuming equal mainstream and coolant specific heats and writing the energy equation for a perfect gas in differential form as

${\text{T}}_{{\text{t}}_{\text{g}}}{\text{W}}_{\text{g}}+{\text{T}}_{{\text{t}}_{\text{c}}}{\text{dW}}_{\text{g}}={(\text{W}}_{\text{g}}+{\text{dW}}_{\text{g}})({\text{T}}_{{\text{t}}_{\text{g}}}+{\text{dT}}_{{\text{t}}_{\text{g}}})$

which, when expanded and second-order terms are dropped, becomes

$\frac{{\text{dT}}_{{\text{t}}_{\text{g}}}}{{\text{T}}_{{\text{t}}_{\text{g}}}}=\frac{{\text{dW}}_{\text{g}}}{{\text{W}}_{\text{g}}}\left(\frac{{\text{T}}_{{\text{t}}_{\text{c}}}}{{\text{T}}_{{\text{t}}_{\text{g}}}}-1\right)$

Since dW/W = ξ, Equation 19.5 may then be written in the more convenient form

$\frac{{\text{ΔP}}_t}{P_{t_g}}=\frac{-\text{γ}}{2}{\text{M}}_{\text{g}}^2\text{ξ}\left[1+\frac{{\text{T}}_{{\text{t}}_{\text{c}}}}{{\text{T}}_{{\text{t}}_{\text{g}}}}-2\frac{{\text{V}}_{\text{c}}}{{\text{V}}_{\text{g}}}{\text{cos ϕ}}_{\text{c}}\right]$ (19.6)

(19.6)

A further simplification results when ${\text{P}}_{{\text{t}}_{\text{c}}}={\text{P}}_{{\text{t}}_{\text{g}}}$, in which

$\frac{{\text{V}}_{\text{c}}}{{\text{V}}_{\text{g}}}=\sqrt{\frac{{\text{T}}_{{\text{t}}_{\text{c}}}}{{\text{T}}_{{\text{t}}_{\text{g}}}}}$

since γ_c = γ_g = γ has been assumed.

The linear approximation method is valuable, not merely because it allows rapid hand calculation, but also since it reveals the primary mixing-loss parameters in their simplest form. This is seen by comparing Equations 19.3 and 19.4 with Equation 19.6. The latter relationship shows that gas properties affect the mixing result only weakly through the specific heat ratio, while mainstream Mach number and coolant fraction are far more significant, as are the coolant-mainstream temperature ratio, velocity ratio, and the coolant injection angle. As would be expected from a one-dimensional analysis, the losses predicted for injection normal to the mainstream flow (ϕ_c = 90°) are independent of coolant velocity since the streamwise component of coolant momentum is zero for all coolant velocities.

Figures 19–24 and 19–25 present a comparison of calculated losses using both of the one-dimensional methods. A constant coolant area for each temperature ratio has been chosen so that coolant and mainstream total pressures are equal at a coolant mass fraction ξ = .03. Coolant velocity for high coolant pressures was limited to the sonic value for these calculations. Aside from showing the close agreement of the linear approximation with constant static pressure mixing, the effects of total temperature ratio and coolant injection angle are quite apparent.

The close agreement of the two calculation methods is not surprising since they are, of course, identical in the limit of zero coolant addition. Figure 19–26 presents the linear approximation in terms of a dimensionless loss parameter in order to show the effects of temperature ratio and coolant injection angle for the ideal case in which coolant and mainstream gases are identical in composition, have equal total pressures, but differ in total temperature. Figure 19–26 is merely a graphic portrayal of the bracketed term in Equation 19.6. The result shows that not only do losses decrease with decreasing angle of injection, but it also reveals that as injection angle decreases, the effect of coolant-to-mainstream temperature ratio becomes less pronounced for higher temperature ratios and more evident for lower ratios. As coolant total temperature approaches extremely low values, the intuition that injection angle becomes less important is confirmed. For total temperature ratios in the general vicinity of ½, however, the effect of coolant injection angle is quite significant.

It is also interesting to note that the one-dimensional methods discussed above are in close agreement with one-dimensional constant area mixing methods, which are far more difficult to apply, particularly for transonic Mach numbers.

General Approach

Attaining the high thrust-to-weight goal will be accomplished by integrating all engineering disciplines. It is estimated that 50% of the goal will come from new materials, 35% from design innovations, and 15% from engine performance (aerodynamic and thermodynamic) advances. Improved engine performance will be helped by higher efficiencies and turbine inlet temperatures. Design innovations will include multifunctional components, high temperature lubricated bearings, 360° static structures, and design simplicity. Fiber-reinforced metal matrix composites (silicon carbide and sapphire fiber), carbon-carbon (C-C), ceramic matrix composite (CMC), intermetallics, and high temperature titanium alloys will be the new materials for future, very advanced, lightweight fighter engines.

The conceptual engine design process begins with the clear enunciation of the requirements. Before the first cycle calculation has been made, the Design Management and Design Team must understand the requirements for performance, weight, cost, and supportability. Requirements for reliability, maintainability, and repairability must be emphasized and given “teeth” by the Design Management.

Engine considerations are initiated with the engine stick flowpath. The “stick” flowpath is a sketch the preliminary engine designer receives from the aerodynamic designers. The engine preliminary designer must know all engine design disciplines and be able to modify the engine flowpath without violating the design intent or engineering laws. The engine preliminary designer at this point must (1) produce a concise compact fan, compressor, combustor, turbine, augmentor, and exhaust nozzle flowpath, (2) mentally and physically determine component mechanical configuration, and (3) determine module configuration. At this point, the preliminary designer must be permitted to struggle alone as he addresses the engine design goals of performance, cost, weight, and supportability.

Once the designer sets design direction involving all design disciplines, the engine’s initial design configuration begins. The skill with which the preliminary designer understands and balances the often-conflicting requirements as he sets the design direction can have a major influence on the supportability characteristics of the final design. If supportability isn’t considered here, it can conceivably be locked out of the final product.

Specific Approach

The crucial first step in the preliminary mechanical design of an engine is the configuration of the sumps and bearings supporting the rotors. Grossly miscalculate initial rotor dynamics and the preliminary design is time wasted. Current engines have multipiece sump, seal, and bearing configurations requiring manual build-up at engine assembly. Future engines will be lubricated by high temperature lubricants with design simplicity as a top priority. Sumps will be one-piece cartridge bearings and seal configurations for easy assembly, maintainability, and removal. The quick replaceability of the cartridge bearing is an example of design simplicity and supportability interaction.

Once the sumps and shafts are in place, remaining engine components can be wrapped around them, starting with the combustor and rotors. Fan, compressor, and turbine rotors are located relative to the bearings as dictated by preliminary rotor dynamic analysis. Advanced compression rotor systems are simple, integrated, lightweight disks and blades called blisks. Compressor blade replaceability is a maintainability tool eliminated by blisks; but to compensate, the rotors will be tolerant of substantial imbalance and will have rugged, damage tolerant, low aspect ratio, high-pressure rise airfoils, and a minimum number of compression stages. Finally, the casings, shells, augmentor, and exhaust nozzle are added to form an initial engine configuration.

Initial Layouts

Once the responsible designer has received and understands the requirements, has initial cycle data in hand, and has held an initial consultation with all design disciplines and management, a series of configuration options should be created by the designer, the designer-draftsperson, and a small team. Modularity and simplicity should be the guidelines. Upon completion, each configuration is critiqued by the full design group and management. These initial layouts will merely be first attempts to incorporate past experience and design practices, but they should also be allowed to contain the new and creative ideas that have been waiting for a home. During this phase of the design process, the drafting room population should be kept to a minimum except when critiques are invited. Supportability interests and considerations, such as modularity, multifunctional components, ruggedness, accessibility, etc., must be given the same emphasis as the other “traditional” requirements during this critique and iteration phase.

Overall Engine

Supportability must be designed into the engine in the preliminary design phase, preferably with little or no increase in engine weight or cost. There must be a game plan for assembly, disassembly, and module selection. Figure 19–33 illustrates a schematic of an engine on an assembly dolly. Here major module assemblies such as the augmentor and exhaust nozzles, fan duct, low pressure turbine, and fan and fan frame are mounted and attached to mini dollies so that after disconnecting electrical and service lines, engine axial disassembly can be attained as illustrated in Figure 19–34. Rapid access to the high-pressure turbine and combustor module is attained.

This type of disassembly/assembly is considered highly desirable in many military scenarios.

The engine core or high pressure section of a turbofan engine, normally a small volume portion of the engine, can be rotated vertically in the dolly for further access to either the high pressure turbine or compressor modules and the compressor rear frame assembly. This engine disassembly approach is nothing new—it has been with us for years. This principle was in General Electric’s J-85 engine design, which now has seen 30 years of service.

Component Design

Figures 19–35 and 19–36 show specific examples of how the maintainability aspect of supportability can be incorporated in the engine in the preliminary design phase. The advent of new composite materials allows the designer to custom tailor materials for specific maintainability problems. Figure 19–35 illustrates a conventional shaft spline detail complete with interference fit pilots. Conventional design practice requires pilot interference fits of 0.0005 inches per inch of pilot radius and special tooling to assemble/disassemble splined components (see Figure 19–37).

This brute force assembly technique has been effective for past rotor designs using monolithic, metal rotor materials, but requires simultaneous heating and chilling or large special tools. However, this technique is antiquated when applied to advanced design rotors of different materials operating at elevated temperatures. By selecting the material of the respective shafts, the thermal growth coefficients can be matched so a loose fit or snug close tolerance pilot fit is obtained at assembly room temperature and a very tight interference fit, caused by differential radial thermal growths, is achieved at pilot/spline operating temperature. This simple, old, but effective technique, where applicable, can be applied throughout the engine, thus eliminating the use of large and heavy special tooling and allowing the use of simple techniques such as “backoff” bolts, as illustrated in Figure 19–35. These simple tools can either remain on the engine or fit in a mechanic’s standard toolbox.

Controlled shaft heating is another feature available when splined shafts have mismatched thermal growths. Selective heating of the inner shaft by a portable electric heater causes inner and outer shaft relative axial growth, unseating the locking nut for easy nut removal. This simple concept, if properly controlled, eliminates the need of heavy load cells or torque multiplier wrenches for rotor assembly/disassembly.

A second simple, but effective, design assembly/disassembly feature is the old technique of “lead-ons” to protect critical and fragile engine sump seals and other components. For decades engine sumps have been designed as an assortment of small components, hand assembled at engine assembly and requiring interim special dimensional checks. More recent engine sump designs have eliminated special assembly dimensional checks but still require delicate assembly procedures. Why not design the bearings, seals, oil jets, and other related components as a one piece mini module? This principle has been accomplished in electronic design—it’s called the circuit board, used extensively in television sets.

If special lead on procedures for components are required, “lead on” rods are a simple and effective method that (1) if left on the engine add little weight, and (2) if removed can fit in a mechanic’s tool box. Long rods or pins, attached via flanges to the engine casing, allow a “lead on” to protect the sumps, shrouds, or other critical components at assembly/disassembly. A circumferential series of rods or pins, with long protrusions, are rigidly attached to one component and engage with and align a second component prior to engagement of the respective component sump seals, rotor shrouds, or other critical parts. A design such as this requires reasonable, but not special, care at assembly.

Another design detail that can be utilized in the preliminary design phase is the old design friend, the curvic coupling. The use of engine compressor blisks (integral one piece disks and blades) reduces compressor rotor weight but can add difficulties to rotor blade repairability; and, in some instances, impairs compressor rotor accessibility by eliminating the compressor split casing. This one piece casing unit can further reduce engine weight. Therefore, another method must be devised for compressor rotor accessibility, and curvic couplings are an example of this.

As stated in Specific Examples—Overall Engine, the engine core module of a turbofan is a small volume and could be rotated vertically in the engine dolly. By machining curvic couplings into the compressor blisk ends, the compressor rotor can be held together as one piece by a single tie bolt. The T700 compressor rotor is a blisk design held together in this manner. Figure 12–36 illustrates the rotor blisk curvic coupling feature. With the engine core module in the vertical position, heat can be applied to the core rotor tie bolt shaft, inducing shaft axial growth and unloading the nut. With the axial load reduced or eliminated, the coupling nut can be removed.

A final example to aid engine supportability in the preliminary design phase is the design of engine static components. Static components in advanced engines account for approximately 60% of the weight. Engine supportability can best be served by minimizing potential problems. There can be no frame problems if there are no frames. Designing 360°, one-piece, multi-functional components that perform functions previously required by two or three units not only saves engine weight but also minimizes supportability problems by component elimination.

In summary, the US military services are participating with the aircraft engine industry in an ambitious program to develop high performance engine technologies for advanced future military applications. This radical program, starting from a “clean sheet of paper,” is a golden opportunity to integrate many radically advanced but fundamentally correct engineering disciplines, including supportability, into future, very high thrust-to-weight fighter engines. The thrust-to-weight ratio goal is double that of current military engines under development and almost triple the thrust-to-weight ratio of fighter engines currently flying. This accomplishment, scheduled for completion in 15 years, will require the total cooperation and understanding between management, supportability, and engineering. Engine supportability techniques may change but fundamental basics will apply. New materials, not yet commercially available, must be understood. The preliminary design engineer will always be a radically thinking, analytically oriented person with his/her head in the clouds; and the supportability engineer will have both feet on the ground beset with day-to-day problems. Some technical crossbreeding between both will be necessary to meet the supportability challenges of the future.