Introduction

The aircraft gas turbine engine is a complex machine using advanced technology from many engineering disciplines such as aerodynamics, materials science, combustion, mechanical design, and manufacturing engineering. In the very early days of gas turbines, the combustor section was frequently the most challenging (Golley, Whittle, and Gunston, 1987). Although the industry’s capability to design combustors has greatly improved, they remain an important design challenge.

This chapter will describe how the combustor interacts with the rest of the engine and flight vehicle by describing the relationship between attributes of the engine and the resulting requirements for the combustor. Emissions, a major engine performance characteristic that relies heavily on combustor design, will be introduced here with more detail found in following chapters. The wide range of operating conditions a combustor must meet as engine thrust varies, which is a major challenge for combustor design, will also be described. Last, the relationship between combustor exit temperature distribution and turbine section durability will be discussed.

When I first became interested in jet engines, smoke trails from the then ultramodern Boeing 707s were an arresting feature of that modern world. Ten years later, smoke was regulated and the U.S. Federal Aviation Administration had canceled the Boeing 2707 supersonic airliner program in the midst of growing environmental concerns. Back in the early 1960s, ground-based gas turbines were a very small business and concern for the environment was only minor. Over the five decades since the 707, the role of gas turbines in our society has greatly expanded, and concern regarding their emissions has grown even faster. Now, the electric power generation gas turbine business has outgrown that of aircraft engines and emissions have become a market discriminator. Indeed, large fortunes have been won and lost on the basis of the emissions performance of land-based gas turbine engines. On the aero engine side, emissions performance is now featured in engine marketing campaigns.

Combustion emissions might be thought an arcane topic. It is certainly complex. It is also of great importance to our society given the dominance of gas turbines for aircraft propulsion and power generation. There are three, basically independent, complicated problems associated with gas turbine emissions – the design of low-emissions combustors, the prediction of the effects of emissions on human health and the global environment, and the formulation of balanced and effective policy and regulation. These challenges are important to three very different groups – technical folk, businesspeople, and policy makers and regulators. This book will be of interest to them all.

Pollution prevention and energy conservation with system efficiency are key elements in arriving at cost-effective long-term solutions that address sustainability to implement national “clean energy” and energy security initiatives. Low air pollution, greenhouse gases, and water impacts are all important to local and regional areas and can be dealt with by some degree of regulatory oversight, with trade-offs appropriately evaluated. International emission standards and regulatory policies for gas turbines described here have developed over the past decade to address some of these challenges.

Gas turbine cogeneration and district energy plants with efficient cycles and reliable dry low NOx combustion can provide important environmental improvements to cleaner energy production. Until recently, GHG emissions and system energy efficiency have not been closely studied in most permitting processes. Pollution prevention planning and environmental assessments may require a more comprehensive strategy, with balanced economic and environmental implementation to allow consideration of a wide range of renewable and cleaner energy choices, including various gas-turbine-based applications.

Definition – Smoke/Soot/Carbonaceous Emissions

Carbonaceous materials emitting from the exhaust of gas turbine engines are frequently referred to as soot emissions, nonvolatile particulates, or smoke. Frequently, these terms are used interchangeably. Such emissions typically consist of single particles ranging from 10–80 nanometers that may agglomerate into a complex fractal chain structure with much larger dimensions. A series of photomicrographs from transmission electron microscope (TEM) analysis at different levels of magnification for a combustor operating at 80 percent power is shown in Figure 5.1.

These carbonaceous soot particles should be contrasted with volatile particulates (see Chapter 6), although volatiles may condense onto soot particles. In particular, soot particles can act as carriers of condensed polycyclic aromatic hydrocarbons (PAHs), some of which are carcinogens. Additional discussion of these subjects can be found in this chapter and in Chapter 6 of this book. Recent results indicate that the morphology of these soot particles may change with power levels and even fuel makeup (Anderson et al., 2011).

Introduction

Premixed combustors for aero engines have been under development for nearly forty years, yet, at the time of writing, the first airplane with premixed combustion still awaits its entry into service. On the other hand, industrial gas turbines have made the transition to premixed combustion within ten years and the level of emissions of nitrogen oxides has decreased tenfold. The differences are due to the peculiarities of gas turbines in flight and a large part of the chapter will be devoted to the understanding of the consequences of those differences for premixed or partially premixed combustion. An obvious difference between both applications lies in the fuels, which are predominantly gaseous for industrial gas turbines and exclusively liquid for aero engines and will continue to be for the foreseeable future. Therefore, premixing in aero combustors always needs to be discussed together with prevaporization, and the differences imposed on the liquid fuel preparation by full or partial prevaporization and premixing are responsible for a large part of the overall development effort. The other determining differences result from the thermodynamic cycles specific to high bypass ratio engines and the impact of the flight profile on the implementation of part load operation. The latter has already been described in Chapter 1.5 and the concept of staging in lean aero engines has been presented in Section 1.5.3 such that this chapter will concentrate more on the implementation of staging and its consequences on the design of the combustor components.

The chapter consists of three parts that partly also follow a historical order: Some results of research are presented that are relevant for lean premixed, prevaporized (LPP) combustion, which for a large part were concurrently achieved with development efforts on LPP combustors. Understanding the limitations and difficulties in the way of fully prevaporized premixed combustion, the concept with the highest emissions reduction potential, will then supply the base for the discussion of partially premixing combustors and their operability aspects.

Introduction

This chapter discusses emissions from systems with extensive levels of exhaust gas recirculation (EGR) or that use oxygen rather than air as a reactant (referred to here as oxyfuel combustion). Such systems have unique attributes that warrant a dedicated chapter in this treatment. First, the systems in which EGR or oxyfuel would be deployed have different degrees of freedom and requirements. For example, both are prominent candidates for carbon capture and storage (CCS) (Griffin et al., 2008; Budzianowski, 2010), where emissions requirements are driven by pipeline or geologic reservoir constraints rather than by atmospheric pollution considerations. Second, while CO2 and H2O dilution have been discussed in Chapters 5 and 7, their presence at very high levels in systems with EGR can provide a significant perturbation of the nominal reactant kinetics (such as in the radical pool) and requires a focused treatment.

As noted earlier, EGR and oxyfuel combustion for gas turbine applications are promising approaches to implement CCS in gas turbine power plants. EGR has also been proposed as a means of promoting fuel flexibility (enabling use of fuels with low heating value (Danon et al., 2010) and high hydrogen content (Lückerath et al., 2008)), and for increasing static stability (resistance to flashback/blowout) (Kalb and Sattelmayer, 2004) and dynamic stability (ElKady et al., 2009) relative to lean premixed combustors, while enabling low levels of pollutant emissions.

The calculation of rotorcraft performance is largely a matter of determining the power required and power available over a range of flight conditions. The power information can then be translated into quantities such as payload, range, ceiling, speed, and climb rate, which define the operational capabilities of the aircraft. The rotor power required is divided into four parts: the induced power, required to produce the rotor thrust; the profile power, required to turn the rotor through the air; the parasite power, required to move the aircraft through the air; and the climb power, required to change the gravitational potential energy. The aircraft has additional contributions to power required, including accessory and transmission losses and perhaps anti-torque power. In hover there is no parasite power, and the induced power is 65% to 75% of the total. As the forward speed increases, the induced power decreases, the profile power increases slightly, and the parasite power increases until it is dominant at high speed. Thus the total power required is high at hover, because of the induced power with a low but reasonable disk loading. At first the total power decreases significantly with increasing speed, as the induced power decreases; then it increases again at high speed, because of the parasite power. Minimum power required occurs roughly in the middle of the helicopter speed range.

The task in rotorcraft performance analysis is the calculation of the rotor forces and power. Procedures to perform these calculations have been developed in the preceding chapters.

Beams and Rotor Blades

An adequate blade structural model is essential for the prediction of rotor loads and stability. Rotor blades almost universally have a high structural fineness ratio and thus are well idealized as beams. The complexities of rotation, and now multiple load paths and composite construction, have required extensive and continuing efforts to develop appropriate beam models for the solution of rotor problems. For exposition of beam theory, particularly relevant to rotor blade analyses, see Hodges (2006) and Bauchau (1985).

A beam is a structure that has small cross-section dimensions relative to an axial line. Based on the slender geometry, beam theory develops a one-dimensional model of the three-dimensional structure. The deflection of the structure is described as functions of the axial coordinate, obtained from ordinary differential equations (in the axial coordinate). The equations depend on cross-section properties, including two-dimensional elastic stiffnesses. The three-dimensional stress field is determined from the deflection variables. Beam theory combines kinematic equations relating strain measures to deflection variables, constitutive equations relating stress resultants to strain measures, and equilibrium equations relating stress resultants to applied loads. When inertial loads are included, the motion is described by partial differential equations, in time and the axial coordinate.

Rotor Configuration

The helicopter rotor type is largely determined by the construction of the blade root and its attachment to the hub. The blade root configuration has a fundamental influence on the blade flap and lag motion and hence on the helicopter handling qualities, vibration, loads, and aeroelastic stability. The basic distinction between rotor types is the presence or absence of flap and lag hinges, and thus whether the blade motion involves rigid-body rotation or bending at the blade root. A simple classification of rotor hubs has the categories articulated, teetering, hingeless, and bearingless, as sketched in Figures 8.1 to 8.4. With real designs (see Figure 1.2) the distinctions are not as clear as in these drawings.

An articulated rotor has its blades attached to the hub with both flap and lag hinges (Figure 8.1). The flap hinge is usually offset from the center of rotation because of mechanical constraints and to improve the helicopter handling qualities. The lag hinge must be offset for the shaft to transmit torque to the rotor. The purpose of the flap and lag hinges is to reduce the root blade loads (since the moments must be zero at the hinge) by allowing blade motion to relieve the bending moments that would otherwise arise at the blade root. With a lag hinge a mechanical lag damper is also needed to avoid a mechanical instability called ground resonance, involving the coupled motion of the rotor lag and hub in-plane displacement.

Hover is the operating state in which the lifting rotor has no velocity relative to the air, either vertical or horizontal. General vertical flight involves axial flow with respect to the rotor. Vertical flight implies axial symmetry of the rotor flow field, so the velocities and loads on the rotor blades are independent of the azimuth position. Axial symmetry greatly simplifies the dynamics and aerodynamics of the helicopter rotor, as is evident when forward flight is considered. The basic analyses of a rotor in axial flow originated in the 19th century with the design of marine propellers and were later applied to airplane propellers. The principal objectives of the analysis of the hovering rotor are to predict the forces generated and power required by the rotating blades and to design the most efficient rotor.

Momentum Theory

Momentum theory applies the basic conservation laws of fluid mechanics (conservation of mass, momentum, and energy) to the rotor and flow as a whole to estimate the rotor performance. The theory is a global analysis, relating the overall flow velocities to the total rotor thrust and power. Momentum theory was developed for marine propellers by W.J.M. Rankine in 1865 and R.E. Froude in 1885, and extended in 1920 by A. Betz to include the rotation of the slipstream; see Glauert (1935) for the history.

Vertical flight of the helicopter rotor at speed V includes the operating states of hover (V = 0), climb (V > 0), and descent (V < 0) and the special case of vertical autorotation (power-off descent). Between the hover and autorotation states, the helicopter is descending at reduced power. Beyond autorotation, the rotor is producing power for the helicopter. The principal subject of this chapter is the induced power of the rotor in vertical flight, including descent. The key physics are associated with the flow states of the rotor in axial flight. Axial flight of a rotor also encompasses the propeller in cruise (V > 0) and static (V = 0) operation, and a horizontal axis wind turbine (V < 0).

Induced Power in Vertical Flight

In Chapter 3, momentum theory was used to estimate the rotor induced power Pi for hover and vertical climb. Momentum theory gives a good power estimate if an empirical factor is included to account for additional induced losses, particularly tip losses and losses due to nonuniform inflow. In the present chapter these results are extended to include vertical descent. Momentum theory is not applicable for a range of descent rates because the assumed wake model is not correct. Indeed, the rotor wake in that range is so complex that no simple model is adequate. In autorotation, the operating state for power-off descent, the rotor is producing thrust with no net power absorption. The energy to produce the thrust (the induced power Pi) and turn the rotor (the profile power Po) comes from the change in gravitational potential energy as the helicopter descends. The range of descent rates where momentum theory is not applicable includes autorotation.