Turbine Engine Compressors PDF

Turbine Engine Compressors Compressor Section The compressor section houses the compressor rotor and works to supply air in sufficient quantity to satisfy the needs of the combustor. Compression results when fuel energy of combustion, and mechanical work of the compressor and turbine,...

Turbine Engine Compressors Compressor Section The compressor section houses the compressor rotor and works to supply air in sufficient quantity to satisfy the needs of the combustor. Compression results when fuel energy of combustion, and mechanical work of the compressor and turbine, are converted into potential energy. The compressor is part of the cold section of the engine. Image by Jeff Dahl licensed under Creative Commons Licence. Adapted for training use by Aviation Australia. Single-spool Axial Flow Turbo-jet Compressors operate on the principle of acceleration of a working fluid followed by diffusion to convert the acquired kinetic energy to a pressure rise. The primary purpose of the compressor is to increase the pressure of the mass of air entering the engine inlet and discharge it to the diffuser, and then to the combustor section, at the correct velocity, temperature and pressure. Compressor efficiency is based on the principle of maximum compression with the least temperature rise. In the compressor, temperature rise is caused by: compression friction. The colder the air is entering the combustor, the greater the rise in temperature during the combustion process. 2022-08-24 B1-15a Gas Turbine Engine Page 78 of 244 CASA Part 66 - Training Materials Only  © NASA Temperature within turbine engine The problems associated with these requirements can be realised if one considers that some compressors must flow air at a velocity of 120–150 m/s (400–500 ft/s) and raise its static pressure perhaps 20–60 times in the space of only a metre or so (a few feet) of engine length. In early compressors, which were less efficient than what we have today, a given amount of work input produced air at a lower pressure and at a higher temperature. To improve on laminar airflow over hundreds of small aerofoils at high velocity and pressure, compressors have been undergoing constant development through the years to achieve optimum efficiency. Presently, this efficiency is said to be in the 80%–90% range. Compressor efficiency is based on the principle of maximum compression with the least temperature rise. Laminar flow minimises friction-induced heat in the air. Secondary purposes of the compressor section are to supply engine service bleed air to cool hot section parts, to pressurise bearing seals, and to supply heated air for inlet anti-icing and fuel system heat for de-icing. Another secondary purpose is to extract air for aircraft uses, and this is usually referred to as customer service bleed air. Common uses for this air include aircraft cabin pressurisation, air conditioning systems, pneumatic starting and various other incidental functions that require clean, pressurised air. The exact location of the bleed ports depends on the pressure or temperature required for a particular job. There are two basic types of compressors: Centrifugal flow Axial flow. Some engines use both types, these are referred to as centri-axial flow. 2022-08-24 B1-15a Gas Turbine Engine Page 79 of 244 CASA Part 66 - Training Materials Only  © Rolls-Royce (1996) The Jet Engine 5th Edition. Adapted for training use by Aviation Australia. Single-entry Two-stage Centrifugal Turbo-propeller Image by Jeff Dahl licensed under Creative Commons Licence. Adapted for training use by Aviation Australia. Axial flow turbojet Relevant Youtube link: Engine Compressor Part 1 (Video) 2022-08-24 B1-15a Gas Turbine Engine Page 80 of 244 CASA Part 66 - Training Materials Only  Centrifugal Flow Compressors Centrifugal Flow Type The centrifugal flow compressor, sometimes referred to as a radial out-flow compressor, is the oldest design and is still in use today. Many smaller engines, as well as the majority of auxiliary power units (APUs) use this design. Advantages are: Light weight Ease of construction Ruggedness. A centrifugal flow compressor assembly consists of: An impeller rotor A diffuser A manifold. Centrifugal compressor components 2022-08-24 B1-15a Gas Turbine Engine Page 81 of 244 CASA Part 66 - Training Materials Only  Impeller The impeller is usually made from aluminium alloy or titanium alloy and can be either single- or dual- sided. The diffuser provides a divergent duct in which the air spreads out, slows down and increases in static pressure. The compressor manifold distributes the air in a turbulence-free condition to the combustion section. The single-sided impeller benefits from ram effect and less turbulent air entry. For this reason, this type of impeller is well suited to many aircraft installations. The single-stage dual-sided impeller design allows for a narrower overall engine diameter and high mass airflow. For this reason, it was favoured in many flight engines in the past. This design does not, however, receive the full benefit from ram effect because the air has to turn radially inwards from a plenum chamber into the centre of the impellers. Impellers can be referred to as single-entry (single-sided) or dual- or double- entry (dual-sided). These terms are use interchangeably throughout this lesson. Impellers 2022-08-24 B1-15a Gas Turbine Engine Page 82 of 244 CASA Part 66 - Training Materials Only  © Rolls-Royce (1996) The Jet Engine 5th Edition. Adapted for training use by Aviation Australia. Double-entry Single-stage Centrifugal Turbo-jet The single-stage dual-sided impeller (pictured above) provided the same compressor ratio and mass air flow with a smaller impeller diameter. Its disadvantages are turbulent side air entry and inefficient ram recovery. It was replaced by single-sided two-stage centrifugal compressors (below). © Rolls-Royce (1996) The Jet Engine 5th Edition. Adapted for training use by Aviation Australia. Single-entry Two-stage Centrifugal Turbo-propeller Impellers incorporate inlet guide vanes called rotating guide vanes, or inducers. They may be part of, or attached to, the impeller. They induce rotation of the air into the eye of the impeller. 2022-08-24 B1-15a Gas Turbine Engine Page 83 of 244 CASA Part 66 - Training Materials Only  Inlet guide vanes are sometimes formed by inclining the front sections of the impeller vanes to impart a whirl motion in the direction of impeller rotation. This eases the change of airflow from the axial to the radial direction. The curved sections may be integral with the radial vanes or formed separately for easier and more accurate manufacture. If inlet guide vanes are not utilised, stationary pre-swirl vanes are situated immediately prior to compressor entry. Rotating guide vanes 2022-08-24 B1-15a Gas Turbine Engine Page 84 of 244 CASA Part 66 - Training Materials Only  Diffuser High-velocity air from the impeller is slung into divergent ducts within the diffuser. The purpose of the diffuser is to convert velocity energy into pressure energy. The diffuser is typically made from one of the following: Aluminium alloy Steel alloy. Diffuser After exiting the diffuser, the high-pressure air enters the manifold. 2022-08-24 B1-15a Gas Turbine Engine Page 85 of 244 CASA Part 66 - Training Materials Only  Manifold The purpose of the manifold is to: Change the airflow direction Deliver air to the combustion chambers. Turning vanes, or cascade vanes, within the manifold straighten the airflow. They are manufactured using aluminium alloy, magnesium alloy or steel alloy. Compressor manifold The compressor manifold distributes the air in a smooth flow to the combustion section. The manifold has one outlet port for each combustion chamber so that the air is evenly divided. A compressor outlet elbow is bolted to each of the outlet ports. The elbows act as air ducts and are often referred to as outlet ducts, outlet elbows or combustion chamber inlet ducts. These outlet ducts change the radial direction of the airflow to an axial direction. To help the elbows perform this function efficiently, turning vanes or cascade vanes are sometimes fitted inside the elbows. These vanes reduce air pressure losses by presenting a smooth turning surface. 2022-08-24 B1-15a Gas Turbine Engine Page 86 of 244 CASA Part 66 - Training Materials Only  Centrifugal Flow Compressor Operation Tip speed of centrifugal impellers reaches approximately Mach 1.3. Radial airflow, however, remains subsonic. The pressure within the compressor casing is capable of preventing airflow separation at low supersonic rotor speeds and causing a high energy transfer to the airflow. The impeller is rotated at high speed by the turbine, and air is continually induced into its eye, or centre. Centrifugal action causes the air to flow radially outwards along the vanes to the impeller tip, thus accelerating the air and causing a pressure rise due to the divergent design of the vanes. The high-velocity air then flows into the diffuser, which fits closely around the periphery of the impeller. There it flows through divergent ducts, where some of the velocity energy is changed into pressure energy. Compressor outputs high pressure air 2022-08-24 B1-15a Gas Turbine Engine Page 87 of 244 CASA Part 66 - Training Materials Only  Diffuser vanes fit closely around the impeller The air, which has slowed down and has had its pressure increased, flows into the manifold through a series of turning vanes. From the manifold, the air flows into the combustion section of the engine. The design of the centrifugal compressor is such that the mass airflow and pressure rise are governed by the rotational speed of the compressor impeller. The most common centrifugal compressor is the single-sided type in either one or two stages. Today the centrifugal compressor is commonly used in conjunction with the axial flow compressor, however it only meets the needs of only smaller flight engines including turboshaft, turboprop and turbofan. All larger engines today are of the axial flow type. Centrifugal types are not used because they impose a serious limitation on mass airflow. Compressor ratios are about the same for single-sided and dual-sided single-stage impellers. Ratios as high as 10:1 can be obtained from a single-stage centrifugal compressor. Compression can be boosted to about 15:1 by a second compressor stage. More than two stages of single entry are impractical because of: Airflow energy loss when making the turns from one impeller to the next High weight per stage High drive power extraction. 2022-08-24 B1-15a Gas Turbine Engine Page 88 of 244 CASA Part 66 - Training Materials Only  © Rolls-Royce (1996) The Jet Engine 5th Edition. Adapted for training use by Aviation Australia. More than two stages of single entry are impractical The centrifugal compressor is shorter in length than an axial compressor, and that is one of its advantages. © Aviation Australia Axial and centrifugal compressors side-by-side 2022-08-24 B1-15a Gas Turbine Engine Page 89 of 244 CASA Part 66 - Training Materials Only  Centrifugal Compressor Multi-spool Some centrifugal flow turboprop engines have two separate rotating assemblies as their gas generator unit: NL compressor and turbine NH compressor and turbine. These are referred to as dual-spool or two-spool engines. At one time, the word spool was used only when describing an axial flow compressor, but many manufacturers today also use it to describe a centrifugal flow compressor. The Pratt and Whitney 100 Series turboprop shown below, for example, has two separately rotating centrifugal compressors plus a power turbine and is described as a three-shaft, two-spool engine. An NL compressor and turbine offer low pressure of low speed. An NH compressor and turbine offer high pressure of high speed. The engine below has a free power turbine to drive the propeller (NP). Dual-spool Centrifugal-flow Free-turbine Turboprop 2022-08-24 B1-15a Gas Turbine Engine Page 90 of 244 CASA Part 66 - Training Materials Only  The advantages of the centrifugal compressor are: High pressure rise per stage, up to 10:1 and 15:1 in a dual stage Good efficiency (compression) from idle to full power Simplicity of manufacture and low cost compared to the axial compressor Low weight overall Low starting power requirements A relatively high FOD tolerance. Disadvantages are: Large frontal area for a given mass airflow More than two stages are not practical and lead to losses in turning the airflow. Centri-Axial Compressors The centrifugal compressor can be used in conjunction with the axial flow compressor, but this typically only meets the needs of smaller flight engines (business jets and helicopters). Although, larger engines today are of the axial flow type, a resurgence of the use of centrifugal compressors has been seen. Recent developments have produced compressor ratios as high as 10:1 from a single centrifugal compressor. Formerly, only axial flow compressors could attain this level of compression. Centri-axial compressor 2022-08-24 B1-15a Gas Turbine Engine Page 91 of 244 CASA Part 66 - Training Materials Only  Axial Flow Compressors Axial Flow Type In an axial compressor, airflow and compression occur parallel to the rotational axis. There are three types of axial flow compressors: single-spool, dual-spool and triple-spool. The single-spool was common in the past for small and large engines, but today it is typically found only in small turboshaft and turboprop engines. The dual-spool is the most common design currently used in large turbofan and turboprop engines. The triple-spool is used in some large and medium- sized turbofan engines. Single Spool In a basic axial flow compressor, the compressor and turbine are connected by a single shaft and rotate as a single unit. Since there is only one compressor unit, the compressor is commonly referred to as a single-spool compressor. While single-spool compressors are relatively simple and inexpensive to manufacture, they do have a few drawbacks. For example, in a long axial compressor, the rear stages operate at a fraction of their capacity, while the forward stages are typically overloaded. Furthermore, the large mass of a single-spool compressor does not respond quickly to abrupt control input changes. Image by Jeff Dahl licensed under Creative Commons Licence. Adapted for training use by Aviation Australia. Single-spool Axial Flow Turbo-jet 2022-08-24 B1-15a Gas Turbine Engine Page 92 of 244 CASA Part 66 - Training Materials Only  Dual Spool The rotors of a dual-spool engine are not mechanically connected, but are linked by an aerodynamic couple. They have two units: Low-pressure (LP or N1) system High-pressure (HP or N2) – gas generator system. The LP compressor boosts compression into the HP compressor. As the HP rotor speed increases, the LP compressor speed increases, but not in direct proportion. The fan is part of the LP system and supplies the first stage of compression. Twin-spool compressor system Dual- and triple-spool axial compressors were developed for the operational flexibility they provide to the engine in the form of high compressor ratios, quick acceleration and better control of stall characteristics. This operational flexibility is not possible with single-spool, axial flow engines. For any given power lever setting, the high-pressure compressor speed is held fairly constant by a fuel control governor. Assuming that a fairly constant energy level is available at the turbine, the low-pressure compressor speeds up and slows down with changes in aircraft inlet conditions resulting from atmospheric changes or flight manoeuvres. The varying low-pressure compressor output therefore provides the high-pressure compressor with the best inlet condition within the limits of its design. That is, the N1 compressor tries to supply the N2 compressor with a fairly constant air pressure for a particular power setting. 2022-08-24 B1-15a Gas Turbine Engine Page 93 of 244 CASA Part 66 - Training Materials Only  The speed of the low-pressure compressor increases with altitude as the atmosphere rarefies from barometric pressure density loss. The increased speed assists in recovering the compressor ratio. As N2 rotor speed increases, N1 increases, but not in direct proportion. In high-bypass engines, the N1 compressor is often referred to as the ‘booster’ stage because it supercharges the N2. Dual spool turbofan engine 2022-08-24 B1-15a Gas Turbine Engine Page 94 of 244 CASA Part 66 - Training Materials Only  Triple Spool The rotors of a triple-spool engine are not mechanically connected. They have two primary units: Fan Gas generator. The fan is the LP system, driven by its own power turbine. The fan supplies the first stage of compression into the gas generator. The gas generator contains the intermediate-pressure (IP) compressor and the high-pressure (HP) compressor. As HP (N3) rotor speed increases, IP (N2) and LP (N1) increase, but not in direct proportion. Note the fan is a separate assembly with its own power turbine and operationally is similar to a free turbine turboprop. The fan is still the first stage of compression. The triple-spool is used in some large and medium-sized turbofan engines. The fan, which is referred to as the LP compressor, the IP compressor and the HP compressor are all driven by separate turbines. The fan turns at a relatively low speed and requires a great deal of torque; therefore, its turbine has multiple stages. Triple spool turbofan engine 2022-08-24 B1-15a Gas Turbine Engine Page 95 of 244 CASA Part 66 - Training Materials Only  Axial Compressor Airflow Control Axial Flow Compressor Stages All axial flow compressors have two main components: rotor stator. A rotor and then a stator make up a compressor stage. Several stages make up the complete compressor. The compressor ratio per stage can increase from 1:1.1 to as much as 1:1.5 (10% to 50%), but depends on the engine type. An axial flow compressor normally has 10–18 stages of compression, increasing the pressure many times more than the intake pressure. Compressor stages 2022-08-24 B1-15a Gas Turbine Engine Page 96 of 244 CASA Part 66 - Training Materials Only  Rotor Blades Each rotor consists of a set of blades fitted into a disc, which move air rearwards through each stage. The speed of the rotor determines the velocity present in each stage and, with increased velocity, kinetic energy is transferred to the air. The stator vanes are placed to the rear of the rotor blades to receive the air at high velocity and act as a diffuser, changing kinetic energy to potential energy (pressure). The stators also have a secondary function of directing airflow to the next stage of compression at the desired angle. Blades fitted to disc Each rotor consists of blades and a disc or, drum. Blades are aerofoils fitted into the disc and move air rearwards through each stage. As explained, rotor rpm determines airflow axial velocity through each stage. Blades are normally made from stainless steel alloy or titanium alloy. Discs are made from nickel steel alloy or titanium alloy. 2022-08-24 B1-15a Gas Turbine Engine Page 97 of 244 CASA Part 66 - Training Materials Only  Blisks Some compressor rotors have one-piece blade and rotor units as shown below. The blades are forged as part of the disc. These one-piece units, called blisks, are commonly used in small turboprop and turboshaft engines. Forged blisk technology is being applied to many smaller fans, compressor rotors and stators, and to some turbine components. Some of the latest large turbofan engines are now being manufactured with multiple stages of blisks, mainly in the HP compressor (GEnx engine, RR Trent XWB, PW1000GTF). Compressor blisk 2022-08-24 B1-15a Gas Turbine Engine Page 98 of 244 CASA Part 66 - Training Materials Only  Compressor Blade Design Compressor blades are constructed with a varying angle of incidence, or twist, similar to that of a propeller. This design feature compensates for the effect on airflow caused by differences in airflow over the different stations of each blade from the base to the tip. The blades also reduce in size from the first stage to the last to accommodate the converging or tapering shape of the compressor housing in which they are rotating. There are several reasons for the shapes of compressor aerofoils. The length, chord, thickness and aspect ratio (ratio of length to width) are calculated to suit the performance factors required for a particular engine and aircraft combination. These are some design aspects common to both compressor and fan blades: A twist is present (called a stagger angle) from base to tip to maintain the exit velocity of airflow at the same value along the blade length. The base area has more camber than the tip area, again to increase axial velocity of airflow and maintain base-to-tip exit velocity. The trailing edge is knife-edge thin to minimise turbulence and provide the best aerodynamic efficiency. Compressor blade design 2022-08-24 B1-15a Gas Turbine Engine Page 99 of 244 CASA Part 66 - Training Materials Only  Blade Root The base or root of a rotor blade often fits loosely into the rotor disc. This loose fit allows for easy assembly and vibration damping. As the compressor rotor rotates, centrifugal force keeps the blades in their correct position, and the airstream over each blade provides a shock-absorbing or cushioning effect. Rotor blade roots are designed in a number of different shapes, such as a bulb, fir tree or dovetail (shown below). To prevent a blade from backing out of its slot, most methods of blade attachment use a pin and a lock tab, locker or snap ring to secure the coupling. Rotor blade root varying designs 2022-08-24 B1-15a Gas Turbine Engine Page 100 of 244 CASA Part 66 - Training Materials Only  Blade Tip The tip of a compressor blade is most important. Some blade tips are squared off, and others have the tip thickness reduced. Those with reduced thickness are called profile tips. The thinner tips have a high natural resonant frequency and are therefore not subject to the vibrations that would affect a blade with a squared tip. The profile tip also provides a more aerodynamically efficient shape for the high-velocity air moved by the blade. These profile tips often touch the housing and make a squealing noise as the engine is shut down. For this reason, profile tips are often called squealer tips. Another blade design that increases compressor efficiency features a localised increase in blade camber, both at the blade tip and blade root. The purpose of this design is to compensate for the friction caused by the boundary layer of air near the compressor case. The increased blade camber helps overcome the friction and makes the blade extremities appear as if they are bent over at each corner, hence the term end bend. Air leakage between profile tips and the compressor housing causes compressor efficiency loss. Loss is prevented by zero running clearance. Zero clearance is obtained by contact and wear between: Profile tip and abradable compressor casing Compressor casing and abradable blade tips. Contact is greatest during run-in. Compressor blades 2022-08-24 B1-15a Gas Turbine Engine Page 101 of 244 CASA Part 66 - Training Materials Only  Stator Vanes Stator vanes are the stationary blades located between each row of rotating blades in an axial flow compressor. As discussed earlier, the stator vanes act as diffusers for the air coming off the rotor, decreasing its velocity and raising its pressure. In addition, the stators help prevent swirling and direct the flow of air coming off each stage to the next stage at the appropriate angle. Like rotor blades, stator vanes have an aerofoil shape. Stator vanes are normally constructed out of stainless steel alloy, nickel steel alloy, steel or nickel because these metals have high fatigue strength. However, titanium may also be used for stator vanes in the low pressure and temperature stages. Stator vanes 2022-08-24 B1-15a Gas Turbine Engine Page 102 of 244 CASA Part 66 - Training Materials Only  Variable Stator Vanes Some stator vanes can be variable in angle in the same manner as inlet guide vanes. A number of stages can be made variable; where an external ring is attached via a lever to an extension of the vane through the casing, the vane’s other end is hinged in its shroud. By rotating the ring slightly, the levers turn the vane to close it as rpm is reduced, thus maintaining airflow directed at the correct angle to the succeeding rotors. Adapted for use by © Aviation Australia Variable stator vanes Stator vanes may be secured directly to the compressor casing or to a stator vane retaining ring, which is secured to the compressor case. Most stator vanes are attached in rows with a dovetail arrangement and project radially towards the rotor axis. Stator vanes are often shrouded at their tips to minimise vibration tendencies. Shrouded stator vanes 2022-08-24 B1-15a Gas Turbine Engine Page 103 of 244 CASA Part 66 - Training Materials Only  Inlet Guide Vanes The set of stator vanes immediately in front of the first-stage rotor blades are called inlet guide vanes. These vanes direct the airflow into the first-stage rotor blades at the best angle while imparting a swirling motion in the direction of engine rotation. This action improves the aerodynamics of the compressor by reducing the drag on the first-stage rotor blades. Some axial compressors with high compressor pressure ratios use variable inlet guide vanes plus several stages of variable stator vanes. These variable inlet guide vanes (IGVs) and stators automatically reposition themselves to maintain proper airflow through the engine under varying operating conditions. Air entering the first stage of the compressor is turned by the IGVs, pictured below. IGVs have a minimum effect on the velocity or pressure. They are mostly fixed, but may be variable in some engines. Air entering the first stage of the compressor is turned by the inlet guide vanes so that it flows in the correct direction to be picked up by the rotor blades. Inlet guide vanes (IGVs) 2022-08-24 B1-15a Gas Turbine Engine Page 104 of 244 CASA Part 66 - Training Materials Only  Axial Flow Compressor Operation Unlike a single centrifugal compressor, which is capable of a compressor pressure ratio of 10:1, a single stage in an axial flow compressor can produce a compressor pressure ratio of approximately 1.25:1 (depending on engine design). Therefore, high compressor pressure ratios are obtained by adding more compressor stages. The task of an axial compressor is to raise air pressure rather than air velocity. Therefore, each compressor stage raises the pressure of the incoming air while the air’s velocity is alternately increased, then decreased as airflow proceeds through the compressor. The rotor blades accelerate the airflow, and then the stator vanes diffuse the air, slowing it and increasing the pressure. The overall result is increased air pressure and relatively constant air velocity from compressor inlet to outlet. As air passes from the front of an axial flow compressor to the rear, the space between the rotor shaft and the stator casing gradually decreases. This shape is necessary to maintain a constant air velocity as air density increases with each stage of compression. To accomplish the convergent shape, each stage of blades and vanes is smaller than the one preceding it. Axial flow compressor operation 2022-08-24 B1-15a Gas Turbine Engine Page 105 of 244 CASA Part 66 - Training Materials Only  Compressor Pressure Ratio The compressor pressure ratio of a gas turbine engine is an extremely important design consideration. In general, the higher the compression, the more remarkable the advantage to the operating cycle in terms of thermal efficiency. In practice, the greater the pressure ratio for a given mass airflow and thrust, the lower the engine fuel consumption. If both pressure ratio and mass increase, thrust increases. But engine weight also goes up as strength of materials is increased. The compressor pressure ratio is determined by measuring the total pressure after the last stage of compression and dividing it by compressor inlet total pressure. Assuming no velocity change between the two points, static pressures could be used to calculate the compressor pressure ratio. Compressor pressure ratio Observe that when compressor inlet total pressure is 14.7 psia and compressor discharge total pressure is 97.0 psia, the compressor pressure ratio is expressed as 97 divided by 14.7, or 6.6 to 1, as indicated in the illustration. The compressor pressure ratio of a compressor is also described in terms of pressure ratio per stage. For example, a business jet may have a small turbofan with an overall compressor ratio of 6.6:1 over eight stages. If we calculate the 8th root of 6.6, we find it to be 1.266, or a 1.266:1 pressure ratio per stage. 2022-08-24 B1-15a Gas Turbine Engine Page 106 of 244 CASA Part 66 - Training Materials Only  For a turbofan engine, the fan supplies the first stage of compression. This boosts the airflow into the LP system of a dual-spool turbofan or into the IP system of a triple-spool turbofan. Compressor Taper As pressure builds in the rear stages of the compressor, velocity tends to drop in accordance with Bernoulli’s principle. This is not desirable because in order to create thrust, the gas turbine engine operates on a principle of velocity change in airflow. Velocity rises and falls through the successive stages of the compressor, but maintains approximately the same inlet and outlet velocity. Even though the pressure is rising dramatically, the velocity is held relatively constant. In order to stabilise the velocity, the shape of the compressor gas path converges, reducing to approximately 25% of the inlet flow area. This tapered shape provides the proper amount of space for the compressed air to occupy. If the compressor blades were all the same length and the air flowed through a constant-area duct, its velocity would decrease as its pressure increased. To keep the air velocity relatively constant as its pressure increases, the rear blades of the compressor are shorter than those at the front, and the passage through which the air flows becomes smaller as the pressure increases. There are two ways to decrease the size of the airflow passage: by holding the outside of the compressor housing constant and increasing the diameter of the drum or discs on which each stage of rotor blades are mounted, or by maintaining the disc or drum diameter and decreasing the outside diameter of the compressor case. Both methods are used. Compressor taper 2022-08-24 B1-15a Gas Turbine Engine Page 107 of 244 CASA Part 66 - Training Materials Only  Cascade Effect So, why does the airflow through an axial flow compressor flow from a low pressure to a higher pressure? The axial compressor is described as containing sets of aerofoils in cascade. This means the aerofoils are arranged in series, which influences air under low pressure in the front stages to flow into an area of higher pressure. The ability of air to flow rearwards against an ever-increasing pressure is similar to forcing water to flow uphill. Pressure must be constantly applied to achieve the correct flow. The idea of constantly applied pressure is explained in the following narrative and drawings. The illustration shows that if a slight positive angle of attack (AOA) exists, a relatively high pressure is present on the bottom of the aerofoil in relation to the pressure on the top of the aerofoil. These high- and low-pressure zones apply to both the rotating aerofoils (rotor blades) and the stationary aerofoils (stator vanes). These zones allow the air in one set of aerofoils to come under the influence of the next set. This is the cascade effect. The illustration depicts high-pressure zone air of the first-stage blade being pumped into the low- pressure zone of its stator. Notice that the stator’s leading edge faces in the opposite direction of the rotor blade’s leading edge, thereby causing the pumping action to occur. The high-pressure zone of the first-stage stator vane then pumps into the low-pressure zone of the second-stage rotor blade. This cascade progress continues through to the last stage of compression. When observing the illustration, it might appear that the rotor blade high- and low-pressure zones could cancel each other out as they blend together, but the overall effect of the divergent shape of the flow path results in a net decrease in velocity and an increase in static pressure. Cascade effect 2022-08-24 B1-15a Gas Turbine Engine Page 108 of 244 CASA Part 66 - Training Materials Only  Compressor Diffuser Section The engine section between the compressor and combustor sections is known as the compressor diffuser because it provides additional space in which air coming from the compressor spreads out. It is a diverging duct and is usually a separate section that is bolted to the rear compressor case. The diffuser case also contains the HP compressor discharge bleed ports and fuel nozzles. The purpose of the diffuser is to reduce the velocity of the air so the flame in the combustion chamber will not be blown out, preparing the air to enter the combustion area. As the air velocity (kinetic energy) decreases, its static pressure (pressure energy) is increased. The diffuser is known as the point of highest pressure in the gas turbine engine. The high wall of pressure it provides, in effect, gives the combustion products something to push against. Compressor diffuser section 2022-08-24 B1-15a Gas Turbine Engine Page 109 of 244 CASA Part 66 - Training Materials Only  Compressor Stall and Surge Compressor aerofoils experience an infinite variety of AOAs and air densities, and controlling the angle of attack (AOA) is a design function of the inlet duct, the compressor and the fuel control sensors. The AOA of the compressor blade is the result of: Inlet air velocity Compressor rpm. The two forces combine to form a vector, the AOA of the aerofoil. Compressor stall is an imbalance between the two vector quantities. A compressor stall is a condition all gas turbine engines experience from time to time, but they are being reduced through FADEC control. When a single compressor blade or stage stalls, it is said to have stalled. When the entire compressor stalls, it is known as surge. One of the characteristics of a gas turbine engine is its tendency to stall under certain operating conditions. Compressor stall occurs in many different types of gas turbine engines. Depending on the operating conditions, stall or surge can occur in various forms and intensities. In their most violent stage, they can cause engine damage and a loud audible noise. Careful inlet designs minimise the chance of an intake-induced stall. AOA of compressor blades As the diagram above shows, the AOA of the compressor blade is determined by the inlet velocity and the rotational speed of the compressor. These two forces combine to form a vector force, which gives us the AOA of the rotor blade aerofoil. A compressor stall is a condition of the airflow when the AOA becomes excessive and the airflow breaks away from the aerofoil. Compressor stalls cause air flowing through the compressor to slow down, to stagnate (stop) or to reverse direction, depending on the stall intensity. 2022-08-24 B1-15a Gas Turbine Engine Page 110 of 244 CASA Part 66 - Training Materials Only  Stall conditions can usually be heard and range in audibility from an air pulsating or fluttering type sound in their mildest form, to a louder pulsating sound, to a sound of violent backfire or explosion. Relevant Youtube link: Boeing Engine Compressor Stall (Video) The remedy for a compressor stall is to reduce the power setting and allow the air inlet velocity and the engine rpm to get back into the correct relationship. Another way to describe stall phenomena in a compressor is by way of a stall or surge margin curve. Every compressor has a best operating point for a particular compressor ratio (Cr), compressor speed (rpm) and mass airflow (Wa), which is commonly called the design point. The normal operating line indicates that the engine will perform without surge or stall at the various compressor pressure ratios, engine speeds and mass airflow along the length of the line, the line falling well below the surge-stall zone. This line represents the maximum Cr and Wa that the compressor is capable of maintaining at a particular rpm. When the three factors are proportionately matched, the engine operates comfortably on the normal operating line. The surge-stall margin is the operating zone between the ‘normal operating line’ and the ‘surge stall line’. The margin decreases as compressor efficiency deteriorates. Relevant Youtube link: Engine Compressors Part 2 (Video) The design point is the point on this line at which the engine operates during most of its service life, that is, cruise speed, at altitude. Stall conditions NOTE: Flames do not often occur during either a surge or a stall – a flame in the inlet would most likely happen with a complete reversal of flow. 2022-08-24 B1-15a Gas Turbine Engine Page 111 of 244 CASA Part 66 - Training Materials Only  Exhaust flames may occur when a stall or surge causes a momentary stagnation of mass airflow followed by an over-rich condition in the combustor. Exhaust flames Compressor Anti-Stall Bleed System The compressor’s ability to pump air is a function of rpm. At low rpm, the compressor does not have the same ability to pump air as it does at a higher rpm. In order to keep the AOA and air velocity within desired limits, it is necessary to unload the compressor in some manner during starting and low power operation. The compressor has less restriction to the flow of air through the use of a compressor surge/stall bleed air valve system. This air is not used for aircraft systems and is dumped directly back into the atmosphere. The pressure within the compressor must be relieved or unloaded. An anti-stall system unloads the compressor by dumping the unwanted air or restricting the inlet airflow during starting and low power operation, and when a pending stall is sensed during any operating condition. The compressor anti-stall bleed system, as with the variable vane system, is installed on some gas turbine engines to minimise compressor acceleration and deceleration stall problems at low and intermediate speeds. Rather than exclude unwanted air, as is the case with the variable vane system, the compressor bleed system automatically dumps away unwanted air. Except at cruise rpm and higher, some compressors cannot handle the amount of air passing through the engine without an air bleed system. 2022-08-24 B1-15a Gas Turbine Engine Page 112 of 244 CASA Part 66 - Training Materials Only  Another way of describing this situation is that in some engines at low and intermediate speeds, a relationship between compressor rotor rpm and airflow cannot be maintained to give the rotating aerofoils the correct effective AOA to the oncoming airstream unless some of the compressor air is being bled away. At high rotational speeds, the compressor is designed to handle maximum airflow without aerodynamic disturbance, so the bleed system is scheduled closed. At the low end of the compressor speed range, the variable vane system allows less air to enter. This in turn keeps compression low and prevents piling up of air molecules in the rear stages, which tends to block airflow. Also at the low end of the compressor speed range, the compressor bleed system bleeds away the excess of air molecules in the rear stages, which in effect accomplishes the very same results. One or more bleed valves fitted to the compressor’s outer case are used to dump unwanted air either into the fan duct or directly overboard. On smaller engines, it is more convenient to use a sliding band which uncovers bleed ports to bleed away unwanted air. On large engines, a combination of bleed valves and variable vanes may be used. The higher the compressor pressure ratio, the greater the need for systems which control the stall margin. Depending on engine type, stall/surge bleed valves can either be a two-positioned, open or closed valve, or variable operation (usually with FADEC control). The bleed valve is fully open when the engine is: Shut down Starting At idle to intermediate power. The bleed valve is fully closed when the engine is: At take-off At cruise power. The bleed band system is incorporated to control the stall margin of the engine. The band is positioned to dump air from a selected rearward stage of compression that will result in the best operating condition of that engine. At low and intermediate speeds, the band is fully open. In the cruise to take-off power range, the band is fully closed. This system does not meter bleed air; it is either fully open or fully closed. 2022-08-24 B1-15a Gas Turbine Engine Page 113 of 244 CASA Part 66 - Training Materials Only  Bleed band system The variable bleed valve (VBV) system lets a part of the LP compressor discharge air to enter the HP compressor. During a fast deceleration, the VBVs prevent LP compressor stall. At low rpm and during reverse, the VBVs open to keep unwanted debris such as water or gravel out of the HP compressor. In general, during steady-state operation, the VBVs close further as N1 increases. The VBVs are closed above 80% N1. The Engine Electronic Controller (EEC) commands the VBVs to be more open during: Rapid deceleration Thrust reverser operation Potential icing conditions. 2022-08-24 B1-15a Gas Turbine Engine Page 114 of 244 CASA Part 66 - Training Materials Only  Summary The bleed-band system is incorporated to control the stall margin of the engine. The band is positioned to dump air from a selected rearward stage of compression that results in the best operating condition of that engine. At low and intermediate speeds, the band is fully open. In the cruise to take-off power range, the band is fully closed. Operation is controlled by the fuel control hydro-mechanical unit (HMU), the EEC or an airflow-sensing transmitter. Variable bleed valves Internal VBVs are modulated open and closed. They open when the engine is: Shut down Starting. At idle to intermediate power, the valves move towards closed as HP (N2) increases, and are fully closed at high power. 2022-08-24 B1-15a Gas Turbine Engine Page 115 of 244 CASA Part 66 - Training Materials Only  Variable bleed valves 2022-08-24 B1-15a Gas Turbine Engine Page 116 of 244 CASA Part 66 - Training Materials Only  Variable Vane System Where high-pressure ratios are required, it becomes necessary to introduce airflow control into the compressor design. This may take the form of variable inlet guide vanes for the first stage plus a number of variable stator vanes for the succeeding stages. As the compressor speed is reduced from its design value, these static vanes are progressively moved towards a preset closed position in order to maintain an acceptable air AOA onto the following rotor blades. At the low end of the compressor speed range, the variable vane system allows less air to enter. This in turn keeps compression low and prevents piling up of air molecules in the rear stages, which tends to block airflow. This system is fuel pressure operated by command of the power lever. It is controlled by fuel signals from the fuel control for its operating schedule. At idle speed, the vanes are scheduled closed, and to provide a stall-free rapid acceleration of the engine, the vanes move towards their open position as engine speed increases. This action maintains the correct AOA relationship between inlet airflow and compressor speed. Variable vane system Relevant Youtube link: VIGVs in Action (Video) 2022-08-24 B1-15a Gas Turbine Engine Page 117 of 244 CASA Part 66 - Training Materials Only  The variable stator vane actuating system is incorporated on many gas turbine engines, especially those with high compression and those in which the compressor may have inherent stall problems during acceleration or deceleration at low or intermediate speeds. The variable vane system automatically varies the geometry (area and shape) of the compressor gas path to exclude unwanted air and maintain the proper relationship between compressor speed and airflow in the front compressor stages. At low compressor speeds, the variable stator vanes are partially closed. As compressor rotor speed increases, the vanes open to allow more and more air to flow through the compressor. In effect, varying the vane angle schedules the correct AOA relationship between the angle of airflow approaching the rotor blades and the rotor blade leading edges. A correct AOA allows for smooth and rapid engine acceleration. Another way of viewing this situation is that the deflection of airflow imposed on the airstream by varying vane angles slows the airstream’s axial velocity before it reaches the rotor blades. Thus the low rpm of the rotor blade and the low axial velocity of the airstream are matched. To control compressor stall and surge, the high-bypass fan engine shown below uses: Variable bleed valves (VBVs) Variable inlet guide vanes (IGVs) Variable stator vanes (VSVs) HPC Bleed Valve (TBV Transient Bleed Valve). Large engine airflow control Relevant Youtube link: Engine Compressors Part 3 (Video) 2022-08-24 B1-15a Gas Turbine Engine Page 118 of 244 CASA Part 66 - Training Materials Only  Axial Flow Advantages There are several advantages of the axial flow compressor: High peak efficiencies (volume) created by its straight-through design. Higher peak efficiencies (compressor pressure ratios) attainable by the use of additional stages of compression. Higher mass airflow for a given frontal area and a low drag coefficient. The disadvantages of the axial flow compressor are: Difficulty and high cost of manufacture Relatively high weight High starting power requirements Low pressure rise per stage (currently around 1.3: 1, but as high as 1.5 to 1.6:1 in newer engines being developed) Good efficiency in the cruise to take-off power range only (poor start, idle and low power operation). The low pressure rise per stage occurs in the axial blade design, where inlet and exit velocities are held at about the same values. 2022-08-24 B1-15a Gas Turbine Engine Page 119 of 244 CASA Part 66 - Training Materials Only  Fan Airflow Control Fan Blades Low-bypass engines have high-aspect-ratio fan blades, normally with mid-span shrouds. They are made from: Aluminium alloy Titanium alloy. A turbofan engine produces thrust similar to that produced by a combination of turbojet and turboprop engines. Some high-bypass engines (those with fan bypass ratios of 4:1 and above) are designed with high- aspect-ratio blades as pictured below. That is, they are long and have a narrow chord. Some of these blades have mid-span shrouds, or snubbers, that form a ring around the fan at the mid-portion of the blade to stiffen it and prevent flutter. The high-bypass-ratio fan blade only became a design possibility with the availability of titanium, conventional designs being machined from solid forgings. A low-weight fan blade is necessary because the front structure of the engine must be able to withstand the large out-of-balance forces that would result from a fan blade failure. Achieving a sufficiently light solid fan blade, even with titanium, required a narrow chord (high aspect ratio). However, with this design, the special feature of a mid-span support (snubber or clapper) is required to prevent aerodynamic instability. This design concept has the disadvantage that the snubber is situated in the supersonic flow where pressure losses are greatest, resulting in inefficiency and a reduction in airflow. © Aviation Australia High aspect ratio fan blades 2022-08-24 B1-15a Gas Turbine Engine Page 120 of 244 CASA Part 66 - Training Materials Only  Low-aspect-ratio (wide chord) blades are coming into wider use today because of their tolerance to foreign objects, especially bird strike damage. In the past, these blades have not been the choice of most designers because of their high weight. Around 1980, hollow titanium blades with composite inner reinforcement materials were developed by Rolls Royce for the RB211. These blades have no stabilising support shrouds and thus produce more mass airflow as a result of the greater flow area. The next advancement was full composite fan blades in the late 1990s by GE on the GE90 engine, followed by swept fan blades on the GE90-115B. Swept fan blades are now becoming the top choice for all new high/ultra-high bypass engines among leading manufacturers. The latest high-bypass engines have low-aspect-ratio fan blades, known as wide chord blades, as shown here. Low-aspect-ratio fan blades Relevant Youtube link: Composite Fan Blades (Video) Normally without mid-span shrouds, they produce a higher mass airflow. Each blade, made from titanium with composite cores or full composite with a metal leading edge sheath, has a predetermined moment weight etched on the root. 2022-08-24 B1-15a Gas Turbine Engine Page 121 of 244 CASA Part 66 - Training Materials Only  Blade with moment weight etched on the root 2022-08-24 B1-15a Gas Turbine Engine Page 122 of 244 CASA Part 66 - Training Materials Only  Fan Blade Replacement With the introduction of the large fan blade, moment weighing of blades has assumed a greater significance. This operation takes into account the mass of each blade and the position of its centre of gravity relative to the centre line of the disc into which it is assembled. The mechanical system of blade moment weighing may be integrated with a computer, which automatically optimises the blade distribution. The moment weight of a blade in grams per millimetre or ounces per inch is identical to the unbalance effect of the blade when installed into a disc. The recorded measurement of blade moment weights enables each blade to be distributed around the disc so that these unbalances are cancelled. On fans with even numbers of blades, blades with similar moment weight are replaced 180° apart. Spare blades are grouped in pairs of similar moment weights. Blade moment weight of two blades does not need to be exactly the same, but must be within the range set out by the manufacturer. The imbalance is corrected by adding different weight balance screws to the fan spinner cone or balance flange on the disc. If a similar-moment-weight blade is not available, replace the damaged blade and the blade 180° opposite. Ground running is necessary to determine if a balance correction will be needed. For fans with even number of blades - Replace fan blade and the fan blade 180° opposite 2022-08-24 B1-15a Gas Turbine Engine Page 123 of 244 CASA Part 66 - Training Materials Only  On fans with odd numbers of blades, blades with similar moment weight are replaced 120° apart. Spare blades are grouped in threes of similar moment weights. Blade moment weight of the three blades does not need to be exactly the same, but must be within the range set out by the manufacturer. The imbalance is corrected by adding different weight balance screws to the fan spinner cone or balance flange on the disc in a similar procedure as for even-blade-number fans. For fans with odd number of blades - Replace fan blade and both the fan blades 120° apart Trim balance is a procedure used to reduce the engine vibration level. It must be performed whenever the engine vibration reaches the level set out in the manufacturer’s Aircraft Maintenance Manual. Usually vibration of this magnitude can be felt in the aircraft cabin. High engine vibration can lead to rapid loss of the engine’s exhaust gas temperature (EGT) margin and engine damage. Engine ground running and fan balance checks are necessary following fan blade replacement, and trim balance procedures may be required if vibration levels approach the limit. This can occur following: Engine deterioration Blending of fan blade damage. 2022-08-24 B1-15a Gas Turbine Engine Page 124 of 244 CASA Part 66 - Training Materials Only  Fan Balancing Balance is corrected by installing heavier or lighter weight screws in the fan spinner. The exact location and weight of the screws must be determined by plotting vector quantities on a polar graph. Always number fan, compressor or turbine blades clockwise (as viewed from the rear) using an approved felt-tip marker. Never use a lead pencil or ballpoint pen. Failure to number the fan blades in the correct direction will cause problems when trying to calculate trim balance. All fan components are index-marked to ensure correct assembly and to maintain balance. Never scratch, punch or etch your own index marks. Damaging highly stressed components in this manner will render them unserviceable as it could lead to catastrophic failure of the component. Fan blade replacement 2022-08-24 B1-15a Gas Turbine Engine Page 125 of 244 CASA Part 66 - Training Materials Only