Aircraft Engines
Piston Engines
From the 1903 flight of the Wright Brothers to the mid-1940's, the piston engine was the only power plant used for aircraft. Originally aircraft engines were built around the same design of car engines. Specifically in the design of cooling systems. Car engines are water or liquid-cooled. They rely on radiators (Heat Exchangers) in order to take air from the outside and cool the warm water coming from the engine (Figure 1).
Figure 1. Schematic of a car engine cooling system.
The problem involving water cooled engines is the excess drag and weight that would be added to the plane thereby having a significant influence in aircraft performance. By 1908 this degradation of aircraft performance due to liquid-cooled systems was noticed and air-cooled engines were first introduced. The savings in weight were substantial. The air-cooled engine weight (on average) was between 30% and 40% of the weight of the liquid-cooled engine.
The first generation of air-cooled engines did not perform as well as expected. In this first generation the cylinders were arranged in a circle around a crankcase. The propeller was fastened to the front of the crankcase and the entire engine rotated! This was done to increase the air velocity over the cylinders for cooling purposes. But what made this type of engine a failure was that the gyroscopic forces were large, limiting the ability of the aircraft to maneuver. This problem was later solved by designing air-cooled engines in a similar radial arrangement, but with the engine fixed and only the propeller rotating. But liquid-cooled engines were not finished just yet, because from about 1915 onward, better designs of liquid-cooled engines were able to develop more power than air-cooled engines. So for the next 25 years, one of the major aircraft design debates concerned the relative drag, weight, and maximum power capabilities of liquid-cooled versus air-cooled engines.
Many of the best fighters of World War II were powered by liquid-cooled engines and by the mid-1940s the debate was over and air-cooled engines were the victors. One major development assisting the air-cooled engines was the development of the NACA cowling. This was an enclosure for the engine that limited the flow of air over the engine cylinders to the air actually in contact with the cooling fins of the cylinders. This advanced was very important in improving the efficiency of air-cooled aircrafts (Figure 2).
Figure 2. Schematic of an improve air-cooled engine with cowling.
From the mid 1940s until today, many advancements have been made in air-cooled engines. The evolution of the propeller-based engines are called turboprops, although the only similar characteristic of today's turboprop and the original air-cooled engines is in that both of them use a propeller in the front.
Air-Breathing Engines
Air breathing engines are also known as gas turbines and aircraft engines are referred to (wrongly) as just turbines. An aircraft gas turbine is a device in which freestream air is taken in through a designed inlet, compressed in a rotating compressor, heated in a combustion chamber, and expanded through a turbine. The gas then exits through a nozzle at a velocity greater than freestream (Figure 3). Though one can classify a piston engine as air breathing, the amount of air taken in by gas turbines is much, much greater. Therefore, we will classify gas turbine engines as air breathing engines.
Figure 3. Schematic diagram of a turbojet engine.
Gas turbines are subdivided into four categories: turboprops, turbofans, turbojets, and prop-fans. The differences are in the speed at which they are used. In reality, the differences in these categories are due to the concepts on which these engines were created. The speed limitations are due to the limitations for the different concepts. The turboprops (Figure 4) are mainly used for cargo planes that fly at speeds typically between 300-500 mph. Its arrangement is derived from that of turbojet engine system. The only difference is that turboprops have a propeller blade and a gear box in front of the engine followed by a similar arrangement as that of a turbojet engine. The turboprop differs from the gas turbine in that the gas turbine drives a propeller. The propeller is the component which generates the thrust for propulsion - not the exhaust jet as in the turbojet case. This is the conceptual difference. The turboshaft is, by the way, in principle also a turboprop.
Figure 4. Schematic diagram of a turboprop engine.
The turbofan (Figure 5) is an engine system similar to the turboprop but with a smaller multiple blade fan encased in a cowling without the gear box. The turbofan is a turbojet with a fan to generate a bypass flow along the core turbojet in order to increase the mass flow. The increased mass flow enhances the efficiency of the engine. The fan is driven by the gas turbine. Note that the fan may be geared or direct driven. This engine system is used for multiple types of aircrafts such as cargo planes, passenger planes, fighter aircrafts, etc. The range of velocities for which this engine is used is typically from below Mach 1.0 (one time the speed of sound) to speeds above Mach 1. Examples of turbofan engines from Pratt and Whitney are the F119-PW-100, F117-PW-100, F100-PW-229, and F100-PW-220 series.
Figure 5. Schematic diagram of a turbofan engine.
The turbojet is the basic, plain, gas turbine. It is a hot gas generator with a jet nozzle to create enough kinetic energy for the propulsion. The turbojet is the most inefficient engine when compared to the turboprop and the turbofan but produces high thrusts. It consumes more fuel than its counterparts but the range of velocities at which is operated cannot be achieved by the turboprop or the turbofan. The top operational velocity is the vicinity of Mach 3.0. Among the most famous aircrafts that make use of this propulsion system are: th eConcorde Anglo-French built , the Russian TU-144 , and the famous American bomber B-52.
The last of the air breathing propulsion systems is the prop-fan (Figure 6). The prop-fan is a newer air breathing system. It is a combination of a turboprop and a turbofan. Its arrangement makes it the only propeller driven aircraft that is capable of pushing an aircrafts to speeds up to Mach one. Its only problem and the main reason for why it has not been incorporated into commercial aircrafts is the intensity of the noise coming from it. It has a higher efficiency than any of the air-breathing engines. It is basically a turbojet engine with a propeller in the back of the engine rather than in front like the turbojet or turbofan.
The propfan need not be a pusher - it can just as well be a tractor, just as the turboprop can be a pusher! The difference between the prop-fan and the turboprop is that in the prop-fan the exhaust jet is used for propulsion in combination with the propeller/fan. Basically, it is a turbofan engine with a higher bypass ratio (ratio of air mass going through bypass to air mass going into engine) and, possibly but not necessarily, without a fan duct. The boundary between the prop-fan and the turbofan is floating, as can be appreciated by different company's vocabulary (e.g., unducted fan engine).
Figure 6. Schematic diagram of a prop-fan engine.
Beyond Turbojets
The turbojet is the general configuration for mostly all aircraft-related propulsion systems. But as flight speed increases beyond Mach 3.5 (3.5 times the speed of sound) the turbojet configuration becomes highly inefficient. The reason for this is that as Mach number increases and the total inlet temperature rises, so does the total inlet pressure. In the vicinity of Mach 3, the inlet pressure rise is sufficient to permit the compressor to be omitted. This would imply that there would be no need for a turbine either since the sole purpose of the turbine is to drive the compressor. The resulting engine is known as a ramjet. A ramjet is probably the simplest and yet most powerful aircraft engine. The ramjet is basically a duct with the front end shaped in the form of an inlet, the aft end designed as a nozzle, and the combustion chamber in the middle (Figure 7). While a ramjet may be operated below Mach 3, they must operate at speeds greater than Mach 3 to be competitive with turbojets.
Figure 7. Schematic diagram of a ramjet.
As speeds increase beyond Mach 5, the temperature at the inlet will increase tremendously. This increase in temperature would tend to dissociate and ionize the air, a process that absorbs energy and reduces the temperature increase sought from the burning of the fuel. In order to correct this problem, a new generation of ramjets was developed known as scramjet or supersonic combustion ramjet. Very few vehicles (mostly missiles) have been fitted with this type of propulsion system.
Aircraft Propulsion
Gas Turbine Operation and Design Requirements
Introduction
A turbine is any kind of spinning device that uses the action of a fluid to produce work. Typical fluids are: air, wind, water, steam and helium. Windmills and hydroelectric dams have used turbine action for decades to turn the core of an electrical generator to produce power for both industrial and residential consumption. Simpler turbines are much older, with the first known appearance dating to the time of ancient Greece.
In the history of energy conversion however, the gas turbine is relatively new. The first practical gas turbine used to generate electricity ran at Neuchatel, Switzerland in 1939, and was developed by the Brown Boveri Company. The first gas turbine powered airplane flight also took place in 1939 in Germany, using the gas turbine developed by Hans P. von Ohain. In England, the 1930's invention and development of the aircraft gas turbine by Frank Whittle resulted in a similar British flight in 1941.
The name "gas turbine'' is somewhat misleading, because to many it implies a turbine engine that uses gas as its fuel. Actually a gas turbine (as shown schematically in Fig. 1) has a compressor to draw in and compress gas (most usually air); a combustor (or burner) to add fuel to heat the compressed air; and a turbine to extract power from the hot air flow. The gas turbine is an internal combustion (IC) engine employing a continuous combustion process. This differs from the intermittent combustion occurring in Diesel and automotive IC engines.
Figure 1. Schematic for a) an aircraft jet engine; and
b) a land-based gas turbine.
Because the 1939 origin of the gas turbine lies simultaneously in the electric power field and in aviation, there have been a profusion of "other names" for the gas turbine. For electrical power generation and marine applications it is generally called a gas turbine, also a combustion turbine (CT), a turboshaft engine, and sometimes a gas turbine engine. For aviation applications it is usually called a jet engine, and various other names depending on the particular engine configuration or application, such as: jet turbine engine; turbojet; turbofan; fanjet; and turboprop or prop jet (if it is used to drive a propeller). The compressor-combustor-turbine part of the gas turbine (Fig. 1) is commonly termed the gas generator
Aircraft Propulsion
Gas Turbine Operation and Design Requirements
Gas Turbine Usage
In an aircraft gas turbine the output of the turbine is used to turn the compressor (which may also have an associated fan or propeller). The hot air flow leaving the turbine is than accelerated into the atmosphere through an exhaust nozzle (Fig. la) to provide thrust or propulsion power:
Figure 1a. Schematic for an aircraft jet engine
Figure 1.b A land-based gas turbine.
A typical jet engine is shown in Fig. 2. Such engines can range from about 100 pounds of thrust (lbst.) to as high as 100,000 lbst. with weights ranging from about 30 to 20,000 lbs. The smallest jets are used for devices such as the cruise missile, the largest for future generations of commercial aircraft. The jet engine of Fig.2 is a turbofan engine, with a large diameter compressor-mounted fan. Thrust is generated both by air passing through the fan (bypass air) and through the gas generator itself. With a large frontal area, the turbofan generates peak thrust at low (takeoff) speeds making it most suitable for commercial aircraft.
Figure 2. A modern jet engine used to power Boeing 777 aircraft.
This is a Pratt & Whitney PW4084 turbofan which can produce 84,000 pounds of thrust. It has a 112-inch diameter front-mounted fan, a length of 192 inches (4.87 m) and a weight of about 15,000 pounds (6804 kg). The nozzle has been disconnected from this engine.
A turbojet does not have a fan and generates all of its thrust from air that passes through the gas generator. Turbojets have smaller frontal areas and generate peak thrusts at high speeds, making them most suitable for fighter aircraft.
In non-aviation gas turbines, part of the turbine power is used to drive the compressor. The remainder, the "useful power", is used as output shaftpower to turn an energy conversion device (Fig. lb) such as an electrical generator or a ship's propeller.
A typical land-based gas turbine is shown in Fig. 3. Such units can range in power output from 0.05 MW(Megawatts) to as high as 240 MW. The unit shown in Fig. 3 is an aeroderivative gas turbine; i.e., a lighter weight unit derived from an aircraft jet engine. Heavier weight units designed specifically for land use are called industrial or frame machines. Although aeroderivative gas turbines are being increasingly used for base load electrical power generation they are most frequently used to drive compressors for natural gas pipelines, power ships and provide peaking and intermittent power for electric utility applications. Peaking power supplements a utility's normal steam turbine or hydroelectric power output during high demand periods ... such as the summer demand for air conditioning in many major cities.
Figure 3. A modern land-based gas turbine used for electrical power production and for mechanical drives. This is a General Electric LM5000 machine with a length of 246 inches (6.2 m) and a weight of about 27,700 pounds (12,500 kg). It produces maximum shaft power of 55.2 MW (74,000 hp) at 3,600 rpm with steam injection. This model shows a direct drive configuration where the l.p. turbine drives both the l.p. compressor and the output shaft. Other models can be made with a power turbine.
Some of the principle advantages of the gas turbine are:
It is capable of producing large amounts of useful power for a relatively small size and weight.
Since motion of all its major components involve pure rotation (i.e. no reciprocating motion as in a piston engine), its mechanical life is long and the corresponding maintenance cost is relatively low.
Although the gas turbine must be started by some external means (a small external motor or other source, such as another gas turbine), it can be brought up to full-load (peak output) conditions in minutes as contrasted to a steam turbine plant whose start up time is measured in hours.
A wide variety of fuels can be utilized. Natural gas is commonly used in land-based gas turbines while light distillate (kerosene-like) oils power aircraft gas turbines. Diesel oil or specially treated residual oils can also be used, as well as combustible gases derived from blast furnaces, refineries and the gasification of solid fuels such as coal, wood chips and bagasse.
The usual working fluid is atmospheric air. As a basic power supply, the gas turbine requires no coolant (e.g. water).
In the past, one of the major disadvantages of the gas turbine was its lower efficiency (hence higher fuel usage) when compared to other IC engines and to steam turbine power plants. However, during the last fifty years, continuous engineering development work has pushed the thermal efficiency (18% for the 1939 Neuchatel gas turbine) to present levels of about 40% for simple cycle operation, and about 55% for combined cycle operation (see next section). Even more fuel-efficient gas turbines are in the planning stages, with simple cycle efficiencies predicted as high as 45-47% and combined cycle machines in the 60% range. These projected values are significantly higher than other prime movers, such as steam power plants.
Aircraft Propulsion
Gas Turbine Operation and Design Requirements
Gas Turbine Cycles
A cycle describes what happens to air as it passes into, through, and out of the gas turbine. The cycle usually describes the relationship between the space occupied by the air in the system (called volume, V) and the pressure (P) it is under. The Brayton cycle (1876), shown in graphic form in Fig. 4a as a pressure-volume diagram, is a representation of the properties of a fixed amount of air as it passes through a gas turbine in operation. These same points are also shown in the engine schematic in Fig. 4b.
Air is compressed from point 1 to point 2. This increases the pressure as the volume of space occupied by the air is reduced. The air is then heated at constant pressure from 2 to 3 in Fig. 4. This heat is added by injecting fuel into the combustor and igniting it on a continuous basis. The hot compressed air at point 3 is then allowed to expand (from point 3 to 4), reducing the pressure and temperature and increasing its volume. In the engine in Fig. 4b, this represents flow through the turbine to point 3' and then flow through the power turbine to point 4 to turn a shaft or a ship's propeller. In an aircraft jet engine, the flow from point 3' to 4 is through the exit nozzle to produce thrust. The "useful work" in Fig. 4a is indicated by the curve 3'- 4. This is the energy available to cause output shaft power for a land-based gas turbine, or thrust for a jet aircraft. The Brayton cycle is completed in Fig. 4 by a process in which the volume of the air is decreased (temperature decrease) as heat is absorbed into the atmosphere.
Figure 4a. Brayton cycle pressure-volume diagram for a unit mass of working fluid (e.g., air), showing work (W) and heat (Q) inputs and outputs.
Figure 4b. Gas turbine schematic showing relative points from the Brayton Cycle diagram.
Most gas turbines operate in an open cycle mode where, for instance, air is taken in from the atmosphere (point 1 in Figs. 4a and 4b) and discharged back into the atmosphere (point 4), with the hot air being cooled naturally after it exits the engine. In a closed cycle gas turbine facility, such as a land-based gas turbine facility, the working fluid (air or other gas) is continuously recycled by cooling the exhaust air (point 4) through a heat exchanger (shown schematically in Fig. 5) and directing it back to the compressor inlet (point 1).
Figure 5. Closed Cycle System.
Because of its confined, fixed amount of gas, the closed cycle gas turbine is not an internal combustion engine. In the closed cycle system, combustion cannot be sustained and the normal combustor is replaced with a second heat exchanger to heat the compressed air before it enters the turbine. The heat is supplied by an external source such as a nuclear reactor, the fluidized bed of a coal combustion process, or some other heat source. Closed cycle systems using gas turbines have been proposed for missions to Mars and other long term space applications.
A gas turbine that is configured and operated to closely follow the Brayton cycle (Fig. 4) is called a simple cycle gas turbine. Most aircraft gas turbines operate in a simple cycle configuration since attention must be paid to engine weight and frontal area. However, in land or marine applications, additional equipment can be added to the simple cycle gas turbine, leading to increases in efficiency and/or the output of a unit. One such modification is reheating.
Reheating occurs in the turbine and is a way to increase turbine work without changing compressor work or melting the materials from which the turbine is constructed. Reheat in a jet engine is accomplished by adding an afterburner at the turbine exhaust, thereby increasing thrust, at the expense of a greatly increased fuel consumption rate.
Aircraft Propulsion
Gas Turbine Operation and Design Requirements
Gas Turbine Components
A greater understanding of the gas turbine and its operation can be gained by considering its three major components (Figs.1, 2 and3 found in the three previous sections): the compressor, the combustor and the turbine. The features and characteristics will be touched on here only briefly.
Compressors and Turbines: The compressor components are connected to the turbine by a shaft in order to allow the turbine to turn the compressor. A single shaft gas turbine fig.(1a and 2b) has only one shaft connecting the compressor and turbine components. A twin spool gas turbine, which is found in land- and marine-based applications, has two concentric shafts, a longer one connecting a low pressure compressor to a low pressure turbine (the low spool) which rotates inside a shorter larger diameter shaft. The shorter, larger diameter shaft connects the high pressure turbine with the higher pressure compressor (the high spool) which rotates at higher speeds than the low spool. A triple spool engine would have a third, intermediate pressure compressor-turbine spool.
Gas turbine compressors are either centrifugal or axial, or can be a combination of both. Centrifugal compressors (with compressed air output around the outer perimeter of the machine) are robust, generally cost less and are limited to pressure ratios of 6 or 7 to 1. They are found in early gas turbines or in modern, smaller gas turbines.
The more efficient, higher capacity axial flow compressors (with compressed air output directed along the center line of the machine) are used in most gas turbines (e.g. Figs. 2 and 3). An axial compressor is made up of a relatively large number of stages, each stage, consisting of a row of rotating blades (airfoils) and a row of stationary blades (stators), arranged so that the air is compressed as it passes through each stage.
Turbines are generally easier to design and operate than compressors, since the hot air flow is expanding rather than being compressed. Axial flow turbines (e.g. Figs. 2 and 3) will require fewer stages than an axial compressor. There are some smaller gas turbines that utilize centrifugal turbines (radial inflow), but most utilize axial turbines.
Turbine design and manufacture is complicated by the need to extend turbine component life in the hot air flow. The problem of ensuring durability is especially critical in the first turbine stage where temperatures are highest. Special materials and elaborate cooling schemes must be used to allow turbine airfoils that melt at 1800-1900°F to survive in air flows with temperatures as high as 3000°F.
Combustors: A successful combustor design must satisfy many requirements and has been a challenge from the earliest gas turbines of Whittle and von Ohain. The relative importance of each requirement varies with the application of the gas turbine, and of course, some requirements are conflicting, requiring design compromises to be made. Most design requirements reflect concerns over engine costs, efficiency, and the environment. The basic design requirements can be classified as follows:
- High combustion efficiency at all operating conditions.
- Low levels of unburned hydrocarbons and carbon monoxide, low oxides of nitrogen at high power and no visible smoke for land-based systems. (Minimized pollutants and emissions.)
- Low pressure drop. Three to four percent is common.
- Combustion must be stable under all operating conditions.
- Consistently reliable ignition must be attained at very low temperatures, and at high altitudes (for aircraft).
- Smooth combustion, with no pulsations or rough burning.
- A low temperature variation for good turbine life requirements.
- Useful life (thousands of hours), particularly for industrial use.
- Multi-fuel use. Characteristically natural gas and diesel fuel are used for industrial applications and kerosene for aircraft.
- Length and diameter compatible with engine envelope (outside dimensions).
- Designed for minimum cost, repair and maintenance.
- Minimum weight (for aircraft applications).
A combustor consists of at least three basic parts: a casing, a flame tube and a fuel injection system. The casing must withstand the cycle pressures and may be a part of the structure of the gas turbine. It encloses a relatively thin-walled flame tube within which combustion takes place, and a fuel injection system.
Compared to other prime movers (such as Diesel and reciprocating automobile engines), gas turbines are considered to produce very low levels of combustion pollution. The gas turbine emissions of major concern are unburned hydrocarbons, carbon monoxide, oxides of nitrogen (NOx) and smoke. While the contribution of jet aircraft to atmospheric pollution is less than 1%, jet aircraft emissions injected directly into the upper troposphere have doubled between the latitudes of 40 to 60 degrees north, increasing ozone by about 20%. In the stratosphere, where supersonic aircraft fly, NOx will deplete ozone. Both effects are harmful, so further NOx reduction in gas turbine operation is a challenge for the 21st century.
Aircraft Engine Thrust Calculations
In this section, we deal with one of the forces acting on an aircraft, namely, the thrust produced by the aircraft's engine. In the first part of this section we will look at propellers and their efficiency. In the second part of this section, we will provide the formula for the thrust of a jet engine.
Total Propeller Efficiency
Propellers are used to drive many lightweight aircraft and were the principal means of propulsion for military aircraft until the advent of the jet engine. As such, it is important to know how propellers work and how efficient they are. The propeller efficiency can never reach the ideal efficiency of 100 %. This is because in the development of the propeller efficiency several concepts are ignored,
1. The friction drag of the blades.
2. The kinetic energy of the rotation of the slipstream.
3. The fact that the thrust is not uniformly distributed over the blades.
The maximun propeller efficiency is about 90 %. This is due to the combined effects of drag from the nacelle and wings upon the propeller. This combined effect drops propeller efficiency to about 87 %. From there the thrust horsepower provided by the propeller is
where:
550 = conversion factor from ft-lbs to horsepower
Thrust Equation For Turbojet-Type Engines
The thrust equation for a turbojet can be derived from the general form of Newton's second law (i.e., force equals the time rate of change of momentum),
The figure below shows the inlet and exhaust flows of the turbojet. The negative thrust due to bringing the freestream air almost to rest just ahead of the engine is called momentum drag or ram drag. The resulting thrust is given by following equation,
Schematic of a turbojet engine.
where: 
