The compressor section of the turbine engine has many functions. Its primary function is to supply enough air to satisfy the requirements of the combustion burners. The compressor must increase the pressure of the mass of air received from the air inlet duct and then discharge it to the burners in the required quantity and pressure.
A secondary function of the compressor is to supply bleed air for various purposes in the engine and aircraft. The bleed air is taken nom any of the various pressure stages of the compressor. The exact location of the bleed port depends on the pressure or temperature required for a particular job. The ports are small openings in the compressor case adjacent to the particular stage from which the air is to be bled. Varying degrees of pressure and heat are available simply by tapping into the appropriate stage. Air often bled from the final or highest pressure stage because at this point pressure and air temperature are at a maximum. At times it may be necessary to cool this high-pressure air. If it is used for cabin pressurization or other purposes where excess heat would be uncomfortable or detrimental the air is sent through a refrigeration unit.
Bleed air has various uses including driving the remote-driven accessories. Some current applications of bleed air are--
Compressor section location depends on the type of compressor. In the centrifugal-flow engine the compressor is between the accessory section and the combustion section; in the axial-flow engine the compressor is between the air inlet duct and the combustion section.
The centrifugal-flow compressor basically consists of an impeller (rotor), a diffuser (stator), and a compressor manifold. The impeller and the diffuser are the two main functional elements. Although the diffuser is a separate component positioned inside and secured to the manifold, the entire assembly (diffuser and manifold) is often referred to as the diffuser.
The impeller's function is to pick up and accelerate air outward to the diffuser. Impellers may be either of two types -- single entry or double entry. Both are similar in construction to the reciprocating engine supercharger impeller. The double-entry type is similar to two back-to- back impellers. However, because of much greater combustion air requirements in turbine engines, these impellers are larger than supercharger impellers.
The principal differences between the two types of impellers are size and ducting arrangement. The double-entry type has a smaller diameter but is usually operated at a higher rotational speed to ensure enough airflow. The single-entry impeller permits convenient ducting directly to the impeller eye (inducer vanes) as opposed to the more complicated ducting necessary to reach the rear side of the double-entry type. Although slightly more efficient in receiving air, the single-entry impeller must be large in diameter to deliver the same quantity of air as the double-entry type. This of course, increases the overall diameter of the engine.
Included in the ducting for double-entry compressor engines is the plenum chamber. This chamber is necessary for a double-entry compressor because air must enter the engine at almost right angles to the engine axis. To give a positive flow, air must surround the engine compressor at a positive pressure before entering the compressor.
Multistage centrifugal compressors consist of two or more single compressors mounted in tandem on the same shaft. The air compressed in the first stage passes to the second stage at its point of entry near the hub. This stage will further compress the air and pass it to the next stage if there is one. The problem with this type of compression is in turning the air as it is passed from one stage to the next.
The diffuser is an annular chamber provide with a number of vanes forming a series of divergent passages into the manifold. The diffuser vanes direct the flow of air from the impeller to the manifold at an angle designed to retain the maximum amount of energy imparted by the impeller. They also deliver the air to the manifold at a velocity and pressure satisfactory for combustion chambers.
The compressor manifold diverts the flow of air from the which, which is an integral part of the manifold, into the combustion chambers. The manifold will have one outlet port for each chamber so that the air is evenly divided. A compressor outlet elbow is bolted to each of the outlet ports. These air outlets are constructed in the form of ducts and are known by a variety of names including "air outlet ducts", "outlet elbows," and "combustion chamber inlet ducts." These outlet ducts perform a very important part of the diffusion process. They change the airflow direction from radial to axial. The diffusion process is completed after the turn. To help the elbows perform this function efficiently, turning vanes (cascade vanes) are sometimes fitted inside the elbows. The vanes reduce air pressure losses by presenting a smooth, turning surface.
The centrifugal compressor is used best on smaller engines where simplicity, flexibility, and ruggedness are primary requirements. These have a small frontal area and can handle high airflows and pressures with low loss of efficiency.
Centrifugal-flow compressors have the following advantages:
They have the following disadvantages:
Axial-flow compressors have two main elements: a rotor (drum or disc type) and a stator. These compressors are constructed of several different materials depending on the load and operating temperature. The drum-type rotor consists of rings that are flanged to fit one against the other so that the entire assembly can be held together by through bolts. This type of construction is satisfactory for low-speed compressors where centrifugal stresses are low (Figure-3-7). The rotor (disc-type) assembly consists of--
Rotor blades are generally machined from stainless steel forgings, although some may be made of titanium in the forward (colder) section of the compressor (Figure 3-8). The blades are attached in the disc rim by different methods using either the fir-tree-type, dovetail-type, or bulb-type root designs. The blades are then locked into place with screws, peening, locking wires, pins, keys, or plates (Figure 3-9). The blades do not have to fit too tightly in the disc because centrifugal force during engine operation causes them to seat. Allowing the blades some movement reduces the vibrational stresses produced by high-velocity airstreams between the blades. The newest advance in technology is a one-piece design machined blade disc (combined disc and blade); both disc and rotor blade are forged and then machined into one (refer to Figure 3-8 again).
Clearances between rotor blades and the outer case are very important to maintain high efficiency. Because of this, some manufacturers use a "wear fit" design between the blade and outer case. Some companies design blades with knife-edge tips that wear away to form their own clearances as they expand from the heat generated by air compression. Other companies coat the inner surface of the compressor case with a soft material (Teflon) that can be worn away without damaging the blade. Rotor discs that are joined together by tie bolts use serration splines or curve coupling teeth to prevent the discs from turning in relation to each other. Another method of joining rotor discs is at their rims.
Axial-flow compressor casings not only support stator vanes and provide the outer wall of the axial paths the air follows but also provide the means for extracting compressor air for various purposes. The stator and compressor cases show great differences in design and construction. Some compressor cases have variable stator vanes as an additional feature. Others (compressor cases) have fixed stators. Stator vanes may be either solid or hollow and mayor may not be connected at their tips by a shroud. The shroud serves two purposes. First, it provides support for the longer stator vanes located in the forward stages of the compressor second, it provides the absolutely necessary air seal between rotating and stationary parts. Some manufacturers use split compressor cases while others favor a weldment, which forms a continuous case. The advantage of the split case is that the compressor and stator blades are readily available for inspection or maintenance. On the other hand the continuous case offers simplicity and strength since it requires no vertical or horizontal parting surface.
Both the case and the rotor are very highly stressed parts. Since the compressor turns at very high speeds the discs must be able to withstand very high certrifugal forces. In addition the blades must resist bending loads and high temperatures. When the compressor is constructed each stage is balanced as a unit. The compressor case in most instances is one of the principal structural, load-bearing members of the engine. It may be constructed of aluminum steel, or magnesium.
Axial-flow compressors have the following advantages:
They have the following disadvantages:
The air in an axial compressor flows in an axial direction through a series of rotating (rotor) blades and stationary (stator) vanes that are concentric with the axis of rotation. Unlike a turbine, which also employs rotor blades and stator vanes the flow path of an axial compressor decreases in cross-sectional area in the direction of flow. This reduces the volume of air as compression progresses from stage to stage.
After being delivered to the face of the compressor by the air inlet duct incoming air passes through the inlet guide vanes. Upon entering the first set of rotating blades, the air, which is flowing in a general axial direction is deflected in the direction of rotation. The air is arrested and turned as it is passed on to a set of stator vanes. Following that it is picked up by another set of rotating blades and soon through the compressor. Air pressure increases each time it passses through a set of rotors and stators.
The rotor blades increase the air velocity. When air velocity increases, the ram pressure of air passing through a rotor stage also increases. This increase in velocity and pressure is somewhat but not entirely nullified by diffusion. When air is forced past the thick sections of the rotor blades static pressure also increases. The larger area at the rear of the blades (due to its airfoil shape) acts as a diffuser.
In the stators velocity decreases while static pressure increases. As air velocity decreases, the pressure due to velocity or ram that has just been gained in Passing through preceding rotor stage decreases somewhat; however, the total pressure is the sum of static pressure and pressure due to ram. Successive increases and decreases in velocity as air leaves the compressor are usually only slightly greater than the velocity of the air at the entrance to the compressor. As the pressure is built up by successive sets of rotors and stators, less and less volume is required. Thus, the volume within the compressor is gradually decreased. At the exit of the compressor, a diffuser section adds the final stage to the compression process by again decreasing velocity and increasing static pressure just before the air enters the engine burner section.
Normally, the temperature change caused by diffusion is not significant by itself. The temperature rise which causes air to get hotter and hotter as it continues through the compressor, is the result of the work being done on the air by the compressor rotors. Heating of the air occurs because of the compression process and because some of the mechanical energy of the rotor is converted to heat energy.
Because airflow in an axial compressor is generally diffusing it is very unstable. High efficiency is maintained only at very small rates of diffusion. Compared to a turbine, quite a number of compressor stages are needed to keep the diffusion rate small through each individual stage. Also, the permissible turning angles of the blades are considerably smaller than those which can be used in turbines. These are the reasons why an axial compressor must have many more stages than the turbine which drives it. In addition, more blades and consequently more stages are needed because the compressor, in contrast to a turbine, is endeavoring to push air in a direction that it does not want to go in.
The dual compressor is a combination either of two axial compressors or of an axial and a centrifugal compressor (Figure 3-10). The dual-axial compressor consists of a low-pressure compressor in front and a high-pressure compressor in the rear. Both compressors (low and high) are driven by two different shafts that connect to different turbines. The starter is usually connected to the high-pressure compressor because it reduces the torque required to start the engine. With the rear (high-pressure) compressor turning at governed Speed, the front (low-pressure) compressor (not governed) is automatically rotated by its turbine. Rotation speed is whatever speed will ensure an optimum flow of air through the compressor. With the front and rear compressor rotors working in harmony instead of interfering with each other, compression rates can be increased without decreasing efficiency. Due to the added length of the engine this type of compressor is found on turbojet aircraft.
Most gas turbines in Army aircraft have a combination of an axial compressor (front) and a centrifugal compressor (rear). The usual combination is a five-or seven-stage axial-flow compressor and a centrifugal-flow compressor. The axial compressor and centrifugal compressor combination is mounted on the same shaft; the compressors turn in the same direction and at the same speed. By combining them, the manufacturer makes the most of the advantages of both compressors small frontal area, increased compression ratios, and shortened overall engine length. Using the centrifugal-flow compressor boosts compression and increases efficiency of the turbine engine. The centrifugal compressor also shortens the length of the engine. If the centrifugal compressor were not added, the manufacturer would have to add more stages of axial compression to equal that of the centrifugal compressor.
The combustion section contains the combustion chambers, igniter plugs, and fuel nozzle or fuel injectors. It is designed to burn a fuel-air mixture and to deliver combusted gases to the turbine at a temperature not exceeding the allowable limit at the turbine inlet. Theoretically, the compressor delivers 100 percent of its air by volume to the combustion chamber. However, the fuel-air mixture has a ratio of l5 parts air to 1 part fuel by weight. Approximately 25 percent of this air is used to attain the desired fuel-air ratio. The remaining 75 percent is used to form an air blanket around the burning gases and to dilute the temperature, which may reach as high as 3500º F, by approximately one-half. This ensures that the turbine section will not be destroyed by excessive heat.
The air used for burning is known as primary air; that used for cording is secondary air. Secondary air is controlled and directed by holes and louvers in the combustion chamber liner. Igniter plugs function during starting only; they are shut off manually or automatically. Combustion is continuous and self-supporting. After engine shutdown or failure to start, a pressure-actuated valve automatically drains any remaining unburned fuel from the combustion chamber. The most common type used in Army gas turbine engines is the external annular reverse-flow type.
The primary function of the combustion section is, of course, to bum the fuel-air mixture, thereby adding heat energy to the air. To do this efficiently, the combustion chamber must--
The location of the combustion section is directly between the compressor and turbine sections. The combustion chambers are always arranged coaxially with the compressor and turbine, regardless of type, since the chambers must be in a through-flow position to function efficiently.
All combustion chambers contain the same basic elements:
There are currently three basic types of combustion chambers, varying in detail only:
The can-type combustion chamber is typical of the type used on both centrifugal and axial-flow engines (Figure 3-11). It is particularly well suited for the centrifugal compressor engine since the air leaving the compressor is already divided into equal portions as it leaves the diffuser vanes. It is then a simple matter to duct the air from the diffuser into the respective combustion chambers arranged radially around the axis of the engine. The number of chambers will vary; in the past as few as 2 and as many as 16 chamber's have been used. The present trend is about 8 or 10 combustion chambers. Figure 3-11 illustrates the arrangement of can-type combustion chambers. On American-built engines these chambers are numbered in a clockwise direction facing the rear of the engine with the No.1 chamber at the top.
Each can-type combustion chamber consists of an outer case or housing with a perforated stainless steel (highly heat-resistant) combustion chamber liner or inner liner (Figure 3-12). The outer case is divided for ease of liner replacement. The larger section or chamber body encases the liner at the exit end; the Smaller chamber cover encases the front or inlet end of the liner.
The interconnector (flame propagation) tubes area necessary part of can-type combustion chambers. Since each can is a separate burner operating independently of the others, there must be some way to spread combustion during the initial starting operation. This is done by interconnecting all the chambers. The flame is started by the spark igniter plugs in two of the lower chambers; it passes through the tubes and ignites the combustible mixture in the adjacent chamber. This continues until all chambers are burning. The flame tubes will vary in construction details from one engine to another although the basic components are almost identical.
The interconnector tubes are shown in Figure 3-12. Bear in mind that not only must the chambers be interconnected by an outer tube (in this case, a ferrule), but there must also be a slightly longer tube inside the outer one to interconnect the chamber liners where the flame is located The outer tubes or jackets around the interconnecting flame tubes not only afford airflow between the chambers but also fulfill an insulating function around the hot flame tubes.
The spark igniters are normally two in number. They are located in two of the can-type combustion chambers.
Another very important requirement in the construction of combustion chambers is providing the means for draining unburned fuel. This drainage prevents gum deposits in the fuel manifold, nozzles, and combustion chambers. These deposits are caused by the residue left when fuel evaporates. If fuel is allowed to accumulate after shutdown there is the danger of afterfire.
If the fuel is not drained, a great possibility exists that at the next starting attempt excess fuel in the combustion chamber will ignite and tailpipe temperature will go beyond safe operating limits.
The liners of can-type combustors have perforations of various sizes and shapes, each hole having a specific purpose and effect on flame propagation in the liner. Air entering the combustion chamber is divided by holes, louvers, and slots into two main streams -- primary and secondary air. Primary (combustion) air is directed inside the liner at the front end where it mixes with the fuel and bums. Secondary (cooling) air passes between the outer casing and the liner and joins the combustion gases through larger holes toward the rear of the liner, cooling the combustion gases from about 3500º F to near 1500º F.
Holes around the fuel nozzle in the dome or inlet end of the can-type combuster liner aid in atomization of the fuel. Louvers are also provided along the axial length of the liners to direct a cooling layer of air along the inside wall of the liner. This layer of air also tends to control the flame pattern by keeping it centered in the liner, preventing burning of the liner walls.
Figure 3-13 illustrates the flow of air through the louvers in the double-annular combustion chamber.
Some provision is always made in the combustion chamber case or in the compressor air outlet elbow for installation of a fuel nozzle. The fuel nozzle delivers the fuel into the liner in a freely atomized spray. The freer the spray, the more rapid and efficient the burning process. Two types of fuel nozzles currently being used in the various types of combustion chambers are the simplex nozzle and the duplex nozzle.
The annular combustion chamber consists basically of a housing and a liner, as does the can type. The liner consists of an undivided circular shroud extending all the way around the outside of the turbine shaft housing. The chamber may be constructed of one or more baskets. If two or more chambers are used, one is placed outside the other in the same radial plane; hence, the term "double-annular chamber."
The spark igniter plugs of the annular combustion chamber are the same basic type used in the can combustion chambers, although construction details may vary. There are usually two plugs mounted on the boss provided on each of the chamber housings. The plugs must be long enough to protrude from the housing into the outer annulus of the double-annular combustion chamber.
The annular-type combustion chamber is used in many engines designed to use the axial-flow compressor. It is also-used by engines incorporating dual-type compressors (combinations of axial flow and centrifugal flow). Its use permits building an engine of small diameter. Instead of individual combustion chambers, the compressed air is introduced into an annular space formed by a combustion chamber liner around the turbine shaft. Usually, enough space is left between the outer liner wall and the combustion chamber housing to permit the flow of cooling air from the compressor. Fuel is introduced through nozzles or injectors connected to a fuel manifold. The nozzle opening may face upstream or downstream to airflow depending on engine design. Various means are provided to introduce primary (compressed) air to the vicinity of the nozzle or injectors to support combustion and additional air dowmstream to increase the mass flow. Secondary cooling air reduces the temperature of gases entering the turbine to the proper level,
Some axial compressor engines have a single annular combustion chamber similar to that shown in Figure 3-14. The liner of this type of burner consists of continuous, circular, inner and outer shrouds around the outside of the compressor drive shaft housing. Holes in the shrouds allow secondary cooling air to enter the center of the combustion chamber. Fuel is introduced through a series of nozzles at the upstream end of the liner. Because of their proximity to the flames all types of burner liners are short-lived in comparison to other engine components; they require more frequent inspection and replacement.
This type of burner uses the limited space available most effectively, permitting better mixing of the fuel and air within a relatively simple structure. An optimum ratio of burner inner surface area to volume is provided; this ensures maximum cooling of the gases as combustion occurs. The design also tends to prevent heat warping However, the burner liner on some engines cannot be disassembled without removing the engine from the aircraft -- a distinct disadvantage.
The latest annular combustion system for military use is a low-pressure fuel injection system with vortex air swirlers to mix fuel and compressor discharge air before combustion. The fuel injector is positioned into the center of an air swirler in the dome of the liner. Fuel leaving the injectors (which has been swirled) is surrounded by a concentric air vortex pattern. This breaks fuel particles down to an extremely small size before they reach the combustion zone, This creates excellent fuel-air mixing that ensures a low smoke level in the exhaust. The low-pressure fuel system does not have tine nozzle orifices and can handle contaminated fuel without clogging.
The can-annular-type combustion chamber was developed by Pratt and Whitney for use in their JT3 axial-flow turbojet engine. Since this engine features the split-spool compressor, it needed a combustion chamber capable of meeting the stringent requirements of maximum strength and limited length plus high overall efficiency. These were necessary because of the high air pressures and velocities in a split-spool compressor along with the shaft length limitations explained below.
The split compressor requires two concentric shafts to join the turbine stages to their respective compressors. The front compressor joined to the rear turbine stages requires the longer shaft. Because this shaft is inside the other, a limitation is imposed on diameter. The distance between the front compressor and the rear turbine must be limited if critical shaft lengths are to be avoided.
Since the compressor and turbine are not susceptible to appreciable shortening the necessary shaft length limitation had to be absorbed by developing a new type of burner. A design was needed that would give the desired performance in much less relative distance than had previously been assigned.
Can-annular combustion chambers are arranged radially around the axis of the engine in this instance the rotor shaft housing. The combustion chambers are enclosed in a removable steel shroud that covers the entire burner section. This feature makes the burners readily available for any required maintenance.
The burners are interconnected by projecting flame tubes. These tubes make the engine-starting process easier. They function identically with those previously discussed but differ in construction details.
Each combustion chamber contains a central bullet-shaped perforated liner. The size and shape of the holes are designed to admit the correct quantity of air at the correct velocity and angle. Cutouts are provided in two of the bottom chambers for installation of the spark igniters. The combustion chambers are supported at the aft end by outlet duct clamps. These clamps secure them to the turbine nozzle assembly.
The forward face of each chamber presents six apertures which align with the six fuel nozzles of the corresponding fuel nozzle duster. These nozzles are the dual-orifice (duplex) type. They require a flow divider (pressurizing valve) as was mentioned above in the can type combustion chamber discussion. Around each nozzle are preswirl vanes for imparting a swirling motion to the fuel spray. This results in better atomization burning and efficiency.
Swirl vanes perform two important functions. They cause--
Swirl vanes greatly aid flame propagation because a high degree of turbulence in the early combustion and cooling stage is desirable. Vigorous mechanical mixing of fuel vapor with primary air is necessary; mixing by diffusion alone is too slow. Mechanical mixing is also done by other measure; for example, placing coarse screens in the diffuser outlet as is done in most axial-flow engines.
Can-annular combustion chambers must also have fuel drain valves in two or more of the bottom chambers. This ensures drainage of residual fuel to prevent its being burned at the next start.
The flow of air through the holes and louvers of the can-annular chambers is almost identical with the flow through other types of burners. Special baffling is used to swirl the combustion airflow and to give it turbulence.
Performance requirements include--
All of the burner requirements must be satisfied over a wide range of operating conditions. For example, airflows may vary as much as 50:1, fuel flows as much as 30:1, and fuel-air ratios as much as 5:1. Burner pressures may cover a ratio of 100:1, while burner inlet temperatures may vary by more than 700º F.
The effect of operating variables on burner performance is--
All turbines in modern jet engines, regardless of the type of compress used, are of of axial-flow design. They consist of one or more stages located immediately to the rear of the engine burner section. Turbines extract kinetic energy from the expanding gases as the gases come from the burners. They convert this energy in to shaft horsepower to drive the compressor and engine accessories. In a turboshaft or turboprop engine one or more turbines will also furnish the power required to turn the engine drive or propeller shaft. Nearly three-fourths of all of the energy available from combustion is needed to drive the compressor or compressors in the case of a dual-compressor engine. This includes the fan of a turbofan engine. If the engine is a turboshaft or turboprop, the turbines are designed to extract as much energy as possible from the gases passing through the engine. So efficient are the turbines in such engines that the propeller in a turboprop aircraft provides approximately 90 percent of the propulsive force with only 10 percent supplied by jet thrust.
The axial-flow turbine has two main elements: turbine rotors (or wheels, as they are sometimes called) and stationary vanes. The stationary part of the assembly consists of a plane of contoured vanes, concentric with the axis of the turbine and set at an angle to form a series of small nozzles. These nozzles discharge the gases onto the blades in the turbine rotors. The stationary vane assembly of each stage in the turbine is usually referred to as the turbine nozzle guide vanes. The turbine nozzle area is the most critical part of the turbine design. If the nozzle area is too large, the turbine will not operate at its best efficiency. If the area is too small the nozzle will have a tendency to choke and lose efficiency under maximum thrust conditions. The turbine nozzle area is defined as the total cross-sectional area of the exhaust gas passages at their narrowest point through the turbine nozzle. It is calculated by measuring and adding the areas between individual nozzle guide vanes.
There are three types of turbines: impluse reaction and a combination of these two known as reaction-impulse. In the impulse type there is no net change in pressure between the rotor inlet and the rotor exit. The blade's relative discharge velocity will be the same as its relative inlet velocity. The nozzle guide vanes are shaped to form passages which increase the velocity and reduce the pressure of the escaping gases. In the reaction type, the nozzle guide vanes do little more in relation to the rotor than alter flow direction. The decrease in pressure and increase in velocity of gases are caused by the convergent shape of the passage between the rotor blades. In a jet engine the turbine is usually a balanced combination of both types known as a reaction-impulse turbine. Its design is intended to achieve both a small diameter and a proper match with the compressor.
Turbines may be either single or multiple stage. When the turbine has more than one stage, stationary vanes are inserted between each rotor wheel and the rotor wheel downstream. They are also placed at the entrance and exit of the turbine unit. Each set of stationary vanes forms a nozzle-vane assembly for the turbine wheel that follows. The exit set of vanes serves to straighten the gas flow before passage through the jet nozzle. The wheels may or may not operate independently of one another, depending on engine type and turbine power requirements.
Shaft RPM, gas flow rate, turbine inlet and outlet temperature and pressure, turbine exhaust velocity, and required power output must all be considered by the designer of the turbine. If the engine is equipped with a dual compressor, the turbine must also be dual or "split." In this event, the forward part of the turbine (which drives the high-pressure compressor) can be single-stage because it receives high-energy gases directly from the burner and turns at a higher RPM than the turbine for the low-pressure compressor. By the time the gases reach the rear part of the turbine (which drives the low-pressure compressor), they have expanded. Considerably more blade area is needed if work or energy balance is to be maintained. To do a multistage turbine is used for the second part of the turbine (Figure 3-15).
Turbines must be designed so that the gases have a high expansion ratio. This results in a large temperature drop in gases passing through the turbine and a cool turbine exhaust. If the engine is equipped with an afterburner, a cool exhaust enables more fuel to be burned in the afterburner without exceeding the temperature limit of the construction materials used in the afterburner.
The turbine wheel is a dynamically balanced unit consisting of super alloy blades attached to a rotating disc. The base of the blade is usually a "fir tree" design to enable it to be firmly attached to the disc and still allow room for expansion. In some turbines the rotating blades are open at their outer perimeter. More commonly, the blade is shrouded at the tip. The shrouded blades form a band around the perimeter of the turbine wheel, which serves to reduce blade vibrations. The weight of the shrouded tips is offset because the shrouds permit thinner, more efficient blade sections than are otherwise possible because of vibration limitations. Also, by acting in the same manner as aircraft wing tip fences, the shrouds improve the airflow characteristics and increase the efficiency of the turbine. The shrouds also serve to cut down gas leakage around the tips of the turbine blades.
Turbines are subjected to high speeds and high temperatures. High speeds result in high centrifugal forces. Turbines must operate close to temperature limits that, if exceeded, lower the strength of the materials they are constructed of. Turbine blades undergo distortion or lengthening known as "creep." Creep means that the blade stretches or elongates. This condition is cumulative. The rate of creep is determined by the load imposed on the turbine and the strength of the blade. The strength of the blade is determined by the temperature within the turbine. Since changes in pitch and creep are more pronounced if engine operating limits are not respected the pilot or flight engineer must closely observe the temperature and RPM limits stipulated by the manufacturer.
The turbine wheel is one of the most highly stressed engine parts. Not only must it operate at temperatures up to approximately 1700º F, but it must do sounder severe centrifugal loads imposed by high rotational speeds of over 40,000 RPM for small engines and 8,000 RPM for larger ones. Consequently, engine speed and turbine inlet temperature must be accurately controlled to keep the turbine within safe operating limits.
The turbine assembly is made of two main parts: the disc and blades. This disc or wheel is a statically and dynamically balanced unit of specially alloyed steel. It usually contains large percentages of chromium, nickel, and cobalt. After forging, the disc is machined all over and carefully inspected using X rays, magnetism, and other inspection methods for structural integrity. The blades or buckets are attached to the disc by means of a fir tree design to allow for different rates of expansion between the disc and the blade while still holding the blade firmly against centrifugal loads. The blade is kept from moving axially either by rivets, special locking tabs or devices, or another turbine stage.
Some turbine blades are open at the outer perimeter (Figure 3-16); in others a shroud is used. The shroud acts to prevent blade-tip losses (gas leakage around the tips of the turbine blade) and excessive vibration. By acting in the same manner as aircraft wing tip fence, the shrouds improve airflow characteristics and increase turbine efficiency.
Shrouds reduce resistance to distortion under high loads, which tend to twist the blade toward low pitch. The shrouded blade has an aerodynamic advantage; thinner blade sections can be used and tip lines can be reduced by using a knife-edge or labyrinth seal at this point.
Shrouding however, requires that the turbine run cooler or at reduced RPM because of the extra mass at the tip. On blades that are not shrouded, the tips are cut or recessed to a knife-edge to permit a rapid "wearing-in" of the blade tip to the turbine casing with am-responding increase in turbine efficiency.
Blades are forged from highly alloyed steel. They are carefully machined and inspected before being certified for use. Many engine manufacturers will stamp a moment weight number on the blade to retain rotor balance when replacement is necessary.
Another method for increasing efficiency is the use of honeycomb shrouding (Figure 3-17). This shroud works as a labyrinth sealing the unshrouded turbine tips. These shrouds are all housed by a stator support, which, in turn, is supported by the engine outer casing. This design is currently in use in the new General Electric turboshaft engines.
Nozzle vanes may be either case or forged. Some vanes are hollow (Figure 3-18) to allow a degree of cooling by compressor bleed air. In all cases the nozzle assembly is made of very high-temperature, high-strength steel to withstand the direct impact of the hot high-pressure, high-velocity gas flowing from the combustion chamber.
Some manufacturers are experimenting with the engine with transpiration-cooled nozzle and turbine blading in which the airflows through thousands of small holes in a porous airfoil made from a sintered wire mesh material (Figure 3-19). The performance of the gas turbine engine depends largely on the temperature at the inlet. Increasing this temperature from the present limit of about 1750o F to the 2500o F possible with transpiration-cooled blades will result in about a 100 percent increase in specific horsepower. Transpiration cooling may be a promising development in gas turbine design.
Design engineers use every device at their command to increase the allowable inlet temperature. On practically all large engines, one such device is to cool the fret-stage turbine inlet guide vanes and the first-stage rotor blades. This is done by conducting compressor bleed air through passages inside the engine to the turbine area. There, the air (the coolant) is led to the longitudinal holes, tubes or cavities in the first-stage vanes and blades.
After entering the vane and blade passages, the air (coolant) is distributed through holes at the leading and trailing edges of the vanes and blades. The air impinges along the vane and blade surfaces and then passes out of the engine with the exhaust. Although bleed air coming from the compressor may be hot, it is cool in relation to the temperature at the turbine inlet. This air, therefore serves to cool the vanes and blades. This permits gases coming from the burner section to enter the turbine at higher temperatures than would otherwise be permissible.
Cooling is necessary only in the turbine inlet area because enough energy is extracted from the exhaust gases by the first or first and second stages of the turbine to reduce the temperature to a tolerable level.
The term "exhaust duet" applies to the engine exhaust pipe or tail pipe including the jet nozzle of a non-after-burning engine (Figure 3-20). Although an afterburner might also be considered a type of exhaust duct, after burning is a subject in itself.
If the engine exhaust gases could be discharged directly to the outside air in an exact axial direction at the turbine exit, an exhaust duct might not be necessary. This, however, is not practical. A larger total thrust can be obtained from the engine if the gases are discharged from the aircraft at a higher velocity than that permissible at the turbine outlet. An exhaust duct is added to collect and straighten the gas flow as it comes from the turbine. It also increases the velocity of the gases before they are discharged from the exhaust nozzle at the rear of the duct. Increasing gas velocity increases its momentum and the thrust produced.
An engine exhaust duct is often referred to as the engine tail pipe. The duct is essentially a simple, stainless steel, conical or cylindrical pipe. The engine tail cone and struts (Figure 3-21) are usually included at the rear of the turbine. The struts support the rear bearing and impart an axial direction to the gas flow, the tail cone helps smooth the flow. Immediately aft of the turbine outlet and usually just forward of the flange to which the exhaust duct is attached, the engine has a sensor for turbine discharge pressure. In large engines, it is not practical to measure internal temperature at the turbine inlet. Therefore, the engine is usually also instrumented for exhaust gas temperature at the turbine outlet. One or more thermocouples preinserted in the exhaust case to provide adequate sampling of exhaust gases. Pressure probes are also inserted in this case to measure pressure of gases coming from the turbine. The gradually diminishing cross-sectional area of a conventional convergent type of exhaust duct is capable of keeping the flow through the duct constant at velocities not exceeding Mach 1.0 at the exhaust nozzle.
Turboshaft engines in helicopters do not develop thrust using the exhaust duct. If thrust were developed by the engine exhaust gas, it would be impossible to maintain a stationary hover; therefore, helicopters use divergent ducts. These ducts reduce gas velocity and dissipate any thrust remaining in the exhaust gases. On find-wing aircraft, the exhaust duct may be the convergent type, which accelerates the remaining gases to produce thrust. This adds additional SHP to the engine rating. Equivalent shaft horsepower (ESHP) is the combination of thrust and SHP.
The rear opening of the exhaust duct is the jet nozzle, or exhaust nozzle as it is often called The nozzle acts as an orifice, the size of which determines velocity of gases as they emerge from the engine. In most non-after-burning engines, this area is critical; for this reason, it is fixed at the time of manufacture. The exhaust (jet) nozzle area should not be altered in the field because any change in the area will change both the engine performance and the exhaust gas temperature. Some early engines, however, were trimmed to their correct RPM or exhaust gas temperature by altering the exhaust-nozzle area. When this is done, small tabs that may be bent as required are provided on the exhaust duct at the nozzle opening. Or small, adjustable pieces called "mice" are fastened as needed around the perimeter of the nozzle to change the area. Occasionally, engines are equipped with variable area nozzles which are opened or closed, usually automatically, with an increase or decrease in fuel flow. The velocity of the gases within a convergent exhaust duct is usually held to a subsonic speed. The velocity at the nozzle approaches Mach 1.0 (the velocity at which the nozzle will choke) on turbojets and low-bypass-ratio turbofans during most operating conditions.
Whenever the pressure ratio across an exhaust nozzle is high enough to produce gas velocities which might exceed Mach 1.0 at the engine exhaust nozzle, more thrust can be gained by using a convergent-divergent type of nozzle (refer back to Figure 3-21). This can be done provided the weight penalty is not so great that the benefit of the additional thrust is nullified. The advantage of a convergent-divergent nozzle (C-D nozzle) is greatest at high Mach numbers because of the resulting higher pressure ratio across the engine nozzle. If the pressure ratio through a subsonic exhaust duct is great enough (this will be the case when the pressure at the entrance to the exhaust duct becomes approximately twice that at the exhaust nozzle), the change in velocity through the duct will be enough to cause sonic velocity (Mach 1.0) at the nozzle. At very high flight Mach numbers, the pressure ratio becomes much more than 20. If a C-D nozzle is used, the velocity at the exhaust nozzle becomes correspondingly greater than Mach 1.0. This is a distinct advantage, provided the nozzle can effectively handle these high velocities.
When a divergent duct is employed in combination with a conventional exhaust duct, it is called a convergent-divergent exhaust duct (Figure 3-21). In the C-D nozzle, the convergent section is designed to handle the gases while they remain subsonic and to deliver them to the throat of the nozzle just as they attain sonic velocity. The divergent section handles the gases after they emerge from the throat and become supersonic further increasing their velocity.
Pressure generated within an engine cannot be converted to velocity, particularly when a convergent nozzle is used. The additional pressure results in additional thrust which, as has been shown, must be added when the total thrust developed by the engine is computed. The additional thrust is developed ineffliciently. It would be much better to convert all of the pressure within the engine to velocity and develop all of the engine thrust by means of changes in momentum. In theory, a C-D nozzle does this. Because it develops this additional part of the total thrust more efficiently, it enables an engine to produce more total net thrust than the same basic engine would generate if it were equipped with a conventional convergent duct and nozzle. The C-D nozzle would be nearly ideal if it could always be operated under the exact conditions for which it was designed. However, if the rate of change in the duct area is either too gradual or too rapid for the calculated increase in weight of the gases, unsteady flow downstream of the throat will occur with an accompanying loss of energy. This ultimately means loss of thrust. If the rate of increase in area of the duct is too little, the maximum gas velocity that can be reached will be limited. If the rate of increase is too great, the gas flow will break away from the surface of the nozzle, and the desired increase in velocity will not be obtained. As exhaust gases accelerate or decelerate with changing engine and flight conditions, their pressure fluctuates above or below the pressure ratio for which the nozzle was designed When this occurs, the nozzle no longer converts all of the pressure to velocity, and the nozzle begins to lose efficiency.
The solution to this dilemma is a C-D nozzle with a variable cross-sectional configuration which can adjust itself to changing pressure conditions. Several types of C-D nozzles have been tried, and a few have been used successfully on production aircraft. As the actual design and operation of such nozzles is usually either classified military information or proprietary information of the manufacturer, the nozzles cannot be described here.
The difficult problem of stopping an aircraft after landing increases many times with the greater gross weights common to large, modern aircraft with their higher wind loadings and increased landing speeds. Wheel brakes alone are no longer the beat way to slow the aircraft immediately after touchdown. The reversible-pitch propeller solved the problem for piston engine and turboprop-powered airplanes. Turbojet and turbofan aircraft, however, must rely on some device such as a parabrake or runway arrester gear or some means of reversing the thrust produced by their engines.
Although sometimes used on military aircraft, the parabrake or drag parachute has distinct disadvantage The parabrake is always subject to either a premature opening or a failure to open at all. The parabrake must be recovered and repacked after each use and, if damaged or lost must be repaired or replaced. Once the parabrake has opened, the pilot has no control over the amount of drag on the aircraft except to release the parachute completely.
Arrester gears are primarily for aircraft carrier deck operation although they are sometimes used by military bases as overshoot barriers for land runways. They would hardly be suitable for commercial airline operation at a busy municipal airport.
The significance of oil system seals in aircraft engines is great. A leaking seal in a turbine engine could cause tire, bearing failure, or cockpit fumes, to name a few dangers. There are three main types of oil system seals: synthetic, labyrinth, and carbon.
Synthetic seals (neoprene, silicone, Teflon, and synthetic rubber) are used throughout the engine's oil system. They are used where metal-to-metal contact would not provide satisfactory sealing to withstand pressures in such items as filters, turbine, and fittings. Seals come in many sizes and shapes and are not normally reused. New replacement seals are received from supply channels usually in a package that prevents damage. In most cases the packages will have a "cure date" stamped on the outside. (Cure date is the date of manufacture of the seals.) This date is particularly important when installing seals retie of rubber, which has a tendency to deteriorate more rapidly than synthetic material. Just as important is to use the proper seal with the correct part number for a specific installation. Never use a seal from another system just because it looks like the right seal. The composition or military specifications may be entirely different which could cause the seal to fail at a crucial moment.
Some synthetic seals coming into contact with synthetic oils such as MIL-7808 or ML-23609, have a tendency to swell; others might deteriorate completely. Occasionally, seals are referred to as "packings" or "gaskets." However, there is a difference between the two. Packing is used to provide a running seal; a gasket is used between two stationary parts to create a static seal. Some manufacturers refer to a gasket as a packing and vise versa. These terms should not be taken literally, Always go strictly by the part number when using seals.
Labyrinth or air seals are designed to allow a small amount of air to flow across the sealing surface. This helps prevent oil (or lower-pressured air) seepage across the same surface (Figure 3-22). Air seals have two separate parts. One part forms a plain or honeycomb surface; the corresponding part is a circular seal with annular grooves. These grooves may use a soft metal as the basic composition or be machined into a surface. Matching the two together (one rotating portion or race with one stationary) forms an air pressure seal. A series of soft metal knifelike edges rides very close to the seal surface or cuts a path into a stationary honeycomb or silver alloy air seal (Figure 3-23).
NOTE: When honeycomb or silver alloy is used, it is bonded to the stationary portion of the air seal.
Air for this seal is normally bled from the compressor and then forced between the sealing surface and the seal. The effect of pressurization prevents oil (or lower-pressured air) from seeping from one section to another during engine operation. Air seals work only when the engine is operating when the engine is shut down seal leakage will occur. Be extremely careful when working in or around the seal area because seals are composed of very soft metal. Avery small nick or groove in a seal may cause a serious oil leak which may require a premature engine change.
Carbon oil seals are used to contain the oil within the bearing areas on most jet turbine engines. All carbon seals form a sealing surface by having a smooth carbon surface rub against a smooth steel surface. The steel surface is called a "seal race" or "faceplate," depending on the engine manufacturer. All carbon seals are preloaded. That is, the carbon must in some way be pressed against the steel surface. Three common preload methods are spring tension, centrifugal force, and air pressure. During operation, the seal may be aided by allowing a small amount of oil to flow into the rubbing surface. The oil also cools the seal as a certain amount of heat is built up by the carbon rubbing on the steel surface.
The carbon oil seal shown in Figure 3-24 consists of two rows of carbon segments (seal ring and back ring) mounted in a housing and held together around their circumference by extension springs. These springs not only hold the segments together by circling the outside but also serve as the preload necessary to press the seals inward. The seal segments nearest the bearing have a lip that forms the seal; the positioning pads contact the steel race and maintain the proper sealing positions of the segments. These positioning pads are sometimes referred to as "wear blocks" because the seal lip is very thin and without the pads would have a short wear life. The grooves between the pads are staggered to reduce airflow toward the sump. On this particular seal, the seal race contact surface is cooled by a spray of oil. The key and spiral pin shown in the figure keep the carbon segments from turning within the seal housing the compression springs press the seal segments into the housing. The entire assembly is held together by the spring retainer and snap ring. This type of seal is stationary and rubs against an inner, rotating seal race.
Other configurations of carbon seals may have several seals on each side of the bearing. They may also rub on the side or outer surface of the seal rather than the inner surface as the one illustrated does. Seals such as the one in Figure 3-25 can be rebuilt by replacing the segments in groups; other seals however, must be returned to an overhaul facility when they are damaged and must be replaced with a complete seal assembly.
High-temperature, high-strength materials and unique methods of manufacture have made the gas turbine engine a practical reality in a few decades. The performance of turbojet and turboprop engines depends largely on the temperature at the inlet to the turbine. Increasing the turbine inlet temperature from the present limit (for most highly produced engines) of approximately 1700o F to 2500o F will result in a specific thrust increase of approximately 130 percent along with a corresponding decrease in specific fuel consumption. For this reason high cycle temperatures are desirable. Not all materials can withstand the hostile operating conditions found in parts of the gas turbine engine.
Metallurgists have been working for almost 50 years improving metals for use in aircraft construction. Each type of metal or alloy has certain properties and characteristics which make it desirable for a particular use, but it may have other qualities which are undesirable. The metallurgist's job is to build up the desirable qualities and tone down the undesirable ones. This is done by the alloying (combining) of metals and by Various heat-treating processes. It is not necessary for the airframe repairer to be a metallurgist, but it is advantageous to have a general knowledge of the properties used in their development. The repairer should be familiar with a few metallurgical terms. The following terms are used in describing the physical properties and characteristics of metals.
In a discussion of metal properties, stress and strain should be mentioned. Stress is a force placed upon a body and is measured in terms of force per unit area, the force being expressed in pounds and the unit of area in square inches; in other words, pounds per square inch (psi). Stress may be in the form of compression, tension, torsion, bending, shearing loads, or a combination of two or more of these. All parts of an aircraft are subjected to stresses. When a part fails to return to its original form after being stressed, it is said to be strained. The various stresses acting on parts of an aircraft, while in flight, have an important bearing on the metals used:
Some of the more commonly used terms in the field of metallurgy are listed below:
Common metal working terms include the following:
The operating conditions inside a gas turbine engine vary considerably, and metals differ in their ability to satisfactorily meet these conditions.
Aluminum Alloys. Aluminum and its alloys are used in temperature ranges up to 500º F. With low density and good strength-to-weight ratios, aluminum forgings and castings are used extensively for centrifugal compressor wheels and housings, air inlet sections, accessory sections, and for the accessories themselves.
Magnesium Alloys. Magnesium is the world's lightest structural metal. Aluminum is 15 times heavier, titanium 25 times heavier, steel 4 times heavier, and copper and nickel alloys are 5 times heavier. Magnesium is combined with small amounts of certain other metals, including aluminum, manganese, zinc, zirconium, thorium, and others, to obtain the strong lightweight alloys needed for structural purposes.
Titanium Alloys. Titanium and its alloys are used for axial-flow compressor wheels, blades, and other forged components in many large, high-performance engines. Titanium combines high strength with low density and is suitable for applications up to 100º F.
Steel Alloys. This group includes high-chromium, molybdenum, high-nickel, and iron-base alloys in addition to low-alloy steels. Because of the relatively low material cost, ease of fabrication, and good mechanical properties, low-alloy steels are commonly used for both rotating and static engine components such as compressor rotor blades, wheels, spacers, stator vanes and structural members. Low-alloy steels can be heat-treated and can withstand temperatures up to 100° F. High nickelchromium iron-base alloys can be used up to 1250º F.
Nickel-Base Alloys. Nickel-base alloys are some of the best metals for use between 1200º F and 1800º F. Most contain little or no iron. They develop high-temperature strength by age hardening and are characterized by long-time creep-rupture strength as well as high ultimate and yield strength combined with good ductility. Many of these materials, originally developed for turbine bucket applications, are also being used in turbine wheels, shafts, spacers, and other parts. Their use is somewhat restricted because of cost and because of their requirement for critical materials.
Cobalt-Base Alloys. Colbalt-base alloys form another important group of high-temperature, high-strength, and high-corrosion-resistant metals. They contain little or no iron. These alloys are used in afterburner and other parts of the engine subjected to very high temperatures.
The number of materials used in alloys is large. Some of the most commonly used elements are listed below.
The percentages of elements used partially determines the physical and chemical characteristics of the alloy and its suitability to a particular application, Tempering and other processes determine the rest. Three characteristics that must be considered are--
High-Temperature Strength. The most highly stressed parts of the gas turbine engine are the turbine blades and discs. Centrifugal forces tending to break the disc vary with the square of the speed. For example, the centrifugal force on a disc rotating at 20,000 RPM will be four times that at 10,000 RPM. Blades weighing only 2 ounces may exert loads of over 4000 pounds at maximum RPM. Blades must also resist the high bending loads applied by the moving gas stream to produce the thousands of horsepower needed to drive the compressor. There is also a severe temperature gradient (difference) of several hundred degrees between the central portion of the disk and its periphery.
Many metals which would be quite satisfactory at room temperatures will lose much of their strength at the elevated temperatures encountered in the engine's hot section. The ultimate tensile strength of a metal at one temperature does not necessarily indicate its ultimate tensile strength at a higher temperature. For example, at l000º F Inconel X has an ultimate tensile strength of approximately 160,000 psi; and S 816 at the same temperature has an ultimate tensile strength of 135,000 psi. At 1500º F their positions are reversed. Inconel X has an untimate tensile strength of 55,000 psi; S 816 has an ultimate tensile strength of 75,000 psi. The creep strength, which is closely associated with ultimate tensile strength, is probably one of the most important considerations in the selection of a suitable metal for turbine blades. Engine vibration and fatigue resistance will also have some influence on the selection and useful life of both discs and blades.
Many materials will withstand the high temperatures encountered in a gas turbine engine (carbon columbium, molybdenum, rhenium, tantalum, and tungsten all have melting points above 4000º F). However, the ability to withstand high temperatures while maintaining reasonable tensile strength is not the only consideration. All of the following qualities must be taken into account when selecting a particular metal:
Resistance to Oxidation and Corrosion. Corrosion and oxidation are results of electrical and chemical reactions with other materials. The hot exhaust gas stream encountered in the engine speeds up this reaction. While all metals will corrode or oxidize, the degree of oxidation is determined by the base alloy and the properties of the oxide coating formed. If the oxide coating is porous or has a coefficient of expansion different from that of the base metal, the base metal will be continually exposed to the oxidizing atmosphere. One solution to oxidation at elevated temperatures is ceramic coatings. Ceramic-coated afterburner liners and combustion chambers are in use today. The ceramic coating has two basic functions:
These coatings are not without disadvantages:
Some promising work is being done with various metal-ceramic combinations called Cermets or Ceramels. Materials being used with ceramics include aluminum, beryllium, thorium, and zirconium oxides, to name a few.
Resistance to Thermal Shock. Many materials which would otherwise be quite suitable must be rejected because of their poor thermal shock characteristics. Several engine failures have been attributed to thermal shock on the turbine disc. Ceramic coating in particular are vulnerable to this form of stress. Improved fuel controls, starting techniques, and engine design have lessened this problem.
The effort to achieve higher turbine inlet temperatures (and therefore higher thermal efficiency) has been approached from two directions: (1) high-temperature materials and (2) cooling methods. A common method of cooling the nozzle guide vanes on gas turbine engines is to pass compressor bleed air through the hollow blades to cool them by convective heat transfer. Some engines also use air bled from the compressor to cool the front and rear face of the turbine discs and the hollow turbine blades.
Transpiration cooling is a novel and efficient method of allowing the turbine blades and other parts within the hot section to operate at much higher turbine inlet temperatures. The Wright Corporation has constructed and run turbine blades at an inlet temperature of 2500º F. In this type of cooled blade the air passes through thousands of holes in a porous airfoil made from a sintered wire mesh material. Since the sintered wire mesh is not strong enough by itself, an internal strut is provided as the main structural support carrying all air foil and centrifugal loads. Fabrication techniques involve rolling layers of woven wire mesh and then sintering these layers to forma porous metal sheet. The sheet is then rolled into an airfoil shape.
Porous materials have been tested for use in combustion chambers and for afterburner liners. A similar material called Rigimesh has also been used in rocket engines to help keep the fuel nozzles cool. Many manufacturers are experimenting with other types of porous materials for use in blades in an attempt to obtain higher turbine inlet temperatures.
Relatively new materials called composites are coming into use in both airframes and engines. In these products graphite, glass or boron filaments are embedded in an epoxy-resin matrix or base substance. Other types of filaments and materials are being tried to meet the demands of higher temperatures and stress The chief advantage of composite material is its very favorable strength-to-weight ratio, which can lead to lighter weight in many structural parts. For example, a lighter fan blade allows a lighter fan disc, which in turn permits a lightening of other parts all the way down the line. Composite materials may be used in conjunction with other load-bearing materials to provide a support function. Typical of this type of structure are fan blades with a steel spar and base and with an airfoil composite shell.
Basic parts of the engine are produced by several casting and forging processes, literally dozens of machine operations, and fabrication procedures using a variety of metal-joining methods.
Casting. Several engine parts are cast in aluminum, magnesium, or steel alloys. These parts include intake and compressor housings, accessory cases, and blading, to name a few. Casting methods differ. They include -
Sand castings. Sand casting uses a wood or metal pattern around which a clay-he sand has been packed to form the mold. The mold is then split, the pattern removed, the mold reassembled, and any cores that are necessary added. Molten metal at a precise temperature is poured into the mold and allowed to cool. The mold is removed, and various heat treatments may be performed to obtain the desired physical characteristics. The casting may be spun while being poured.
Spin casting. Spin casting results in a denser, sounder casting. Spinning is normally performed on small ring sections. Cooling of the metal radially inward results in fewer stresses.
Lost-wax or investment casting. Basically, the investment casting process involves the use of heat-disposable wax or plastic patterns which are surrounded with a refractory material to form a monolithic mold Patterns are removed from the mold in ovens, and molten metal is poured into the hot mold. Sometimes this pouring is done in a vacuum furnace. After cooling the mold material is quite fragile and easily removed from the casings. Because the finished product duplicates the pattern exactly, the fabrication of patterns is critical. They are made by injecting molten wax or plastic into metal dies. The finished castings have an exceptionally smooth surface finishing and require very little further machining. Incidentally, this process is not new. It was used by the ancient Greeks and Egyptians to cast lightweight statues, intricate bowls, and pitchers.
Resin-shell mold casting. Resin-shell mold casting is a high-production method similar to investment casting except that the tolerances are not held as closely. In many ways it rivals sand casting in economy.
Slip casting. Slip casting (which was borrowed from the ceramics industry) is used for super heat-resistant materials. Often it is the only way certain materials can be shaped. Metal ceramics, silicon nitride, and refractory metals cast this way can withstand temperatures over 2200º F.
Mercasting. The Mercast process is a precision casting technique. It is essentially the same kind of method as the lost-waxer investment process except that frozen mercury is used as a pattern instead of wax. Liquid mercury is poured into a master mold where it is frozen at temperatures below 40º F. Then it is removed and created with a cold refractory slurry to a thickness of 1/8 inch or more. The refractory shell is dried at low temperature then the shell and mercury are brought to room temperature and the mercury is melted out. The refractory shell is fired to give it strength and then is used as the mold for a usual casting process. Complicated parts can be made with the Mercast process. Very close tolerances and excellent surface finish can be obtained. The cost, however, is higher than that of some other methods.
Forging. Disks, drive Shafts, gears, vanes, blades, and numerous other parts of the gas turbine engine are manufactured by forging. This process allows the development of a grain structure and results in a finely grained, more ductile, strong dense product. Forging is by rapid hammering or slow pressing. The choice of technique depends on the resistance of metal to rapid deformation. The workpiece is generally heated to improve plasticity and reduce forging forces. It will often pass through several different dies before the final shape is obtained. All ductile materials can be forged but their forgeability varies considerably. Forgeability generally depends upon--
Machining. Common tools used to manufacture gas turbine parts include lathes, mills, broaches, grinders, shapers and planers, polishers and buffers, drills, saws, shears, filers, threaders, contour machines of all kinds, and a host of other devices to cut and form metal. Many of these devices use a numerical tape control or other automatic control devices to reduce human error and produce a more uniform, less expensive product. Robots equipped with computers also assist in machining parts.
Some nontraditional machining techniques for removing metal from super hard and super tough alloys and other materials whose complex shapes preclude conventional metal-cutting tools include--
Other nonconventional machining includes everything from abrasive jet cutting to ultrasonic machining.
Chemical milling. Chemical milling involves the removal of metal by dissolving it in a suitable chemical. Those areas that are not to be dissolved are masked with nonreactive materials. The process can be used on most metals, including aluminum, magnesium, titanium, steels, and superalloy for surface sculpturing. Both sides of the workpiece can be chemically milled simultaneously. In addition, the process can be used to machine very thin sheets.
Electrochemical machining. Electrochemical machining is basically a chemical deplating process in which metal, removed from a positively charged workpiece using high-amperage, low-voltage DC, is flushed away by a highly pressurized electrolyte before it can plate out on the cathode tool. The cathode tool is made to produce the desired shape in the workpiece; both must be electrically conductive. The work proceeds while the cathode and workpiece are both submerged in an electrolyte such as sodium chloride. A variation and extension of electrochemical machining is electro-stream drilling. In this process a negatively charged electrolyte, usually an acid, drills holes in a workpiece that has been positively charged Holes as small as 0.005 inch in diameter and 0.5 inch deep in superalloys can be drilled in this manner.
Electric-discharge machining. In electric-discharge machining high voltages are used to produce a high electrical potential between two conductive surfaces (the workpiece and electrode tool) both of which are immersed in a dielectric fluid. Material is removed from both the electrode and the workpiece by a series of very short electric discharges or sparks between the two and is swept away by the dielectric fluid. More material is removed from the workpiece than from the tool by proper selection of the two materials. This process can be used to shape complex parts to very close tolerances from refractory metals and alloys that were formerly impossible to machine. The use of electric-discharge machining is limited in that it is slower than electrochemical machining tool replacement can become expensive, and the surface of the workpiece is damaged as a result of the sparks. On the other hand, the EDM process is less expensive than the ECM process.
Electron-and laser-beam machining. Electron-beam machining and laser-beam machining are being used experimentally. They may find future use in the production of gas turbines and other aerospace components.
Fabrication. Welding is used extensively to fabricate and repair many engine parts. Fabricated sheet steel is used for combustion chambers, exhaust ducts, compressor casings, thrust reversers, silencers, and so forth. Common methods include resistance and inert-gas (usually argon) welding. Less common methods use plasmas and lasers. Electric resistance welding is used to make spot, stitch (overlapping spots), and continuous-seam welds. Inert-gas welding employs a nonconsumable electrode (tungsten-thorium alloy) surrounded by some inert gas such as argon or helium. The gas prevents an adverse reaction with the oxygen present in the normal atmosphere. The inert gas can be applied in the immediate area of the arc. In the case of production runs the workpiece or the entire welding machine can be enclosed in a thin plastic balloon, sometimes as large as a room. The entire plastic bubble is filled with and supported by the inert gas. The operator stands on the outside and works through specially designed armholes. After welding, many parts must be stress-relieved. Where temperature or working loads are not large, brazing or silver soldering may be used to join such parts as fittings and tube assemblies.
Electron-beam welding is showing great promise as a method of fabricating parts from heretofore difficult to weld or unweldable materials. Electron-beam welding uses a stream of focused electrons traveling at speeds approaching 60 percent of the speed of light. Even though the mass of electrons which form the beam is small they are traveling at such speeds that they contain a great amount of kinetic energy. When the beam strikes the workpiece, the kinetic energy is transformed into heat energy. The welding usually takes place in a vacuum, although nonvacuum techniques can be used. The following characteristics make this welding process a valuable one in the gas turbine manufacturing area:
Finishing. The basic material, the properties desired in the finished product, and the kind of protection desired determine the type of surface and internal treatment received. The variety is considerable and includes the following:
The Coating Service of Union carbide Corporation has developed and is producing machines for applying extremely wear-resistant and other specialized coating to gas turbine parts, tools, and other machines. The different coatings are applied by either of two methods- the detonation gun (D-gun) or the plasma gun. Four times a second, a spark ignites the mixture and creates a detonation which hurls the coating particles, heated to a plastic state by the 6000º F temperature in the gun, out of the barrel at a speed of 2500 feet per second. The part to be plated is kept below 300º F by auxiliary cooling streams. The high-level sound of 150 decibels necessitates housing the gun in a double-walled, sound-insulated construction. Operation is controlled from outside this enclosure.
The plasma gun or torch produces and controls a high-velocity, inert-gas stream that can be maintained at temperatures above 20,000º F. Unlike the D-gun process no combustion takes place. The high-temperature plasma is formed by ionizing argon gas in the extreme heat of an electric arc. Gas molecules absorb heat energy from the arc split into atoms, and then further decompose into electrically charged particles called ions. The hot gas stream can melt any known material, without decomposition. When the molten particles, which are introduced in powdered form, strike the part being coated, a permanent welded bond is formed. While the D-gun is a patented Union Carbide machine, other manufacturers make and distribute a variety of plasma-plating and cutting devices.
Heat Treatments. All of the following procedures alter the mechanical properties of steel to suit the end
Teflon, nylon, carbon, rubber, Bakelite, and a host of plastic materials are used in the gas turbine engine mainly as sealing and insulation materials. For example, nylon and Teflon are used to insulate and protect the shielded electrical wiring located on the outside of the engine. Teflon is also used on the J-79 F4 Phantom engine for the sealing the variable-stator-vane actuators. Carbon is used largely inside the engine in the form of carbon rubbing seals. Some of these "face" carbon rubbing seals must be flat to within two helium light bands, or approximately 23 millionths of an inch. Rubber and rubberized fabric materials makeup the sealing edge of the fire seal which divides the hot and cold sections of the engine when mounted in the nacelle. Synthetic rubber is used extensively throughout the engine in the form of O-rings or other shaped seals.