Saturday, June 20, 2015
FREE AVIATION STUDY: BIOGRAPHY OFATYPICAL ENGINE
FREE AVIATION STUDY: BIOGRAPHY OFATYPICAL ENGINE: BIOGRAPHY OFATYPICAL ENGINE This concluding section of the chapter is presented to show that all specialties have a long way to go befor...
BIOGRAPHY OFATYPICAL ENGINE
BIOGRAPHY OFATYPICAL ENGINE
This concluding section of the chapter is presented to show that all
specialties have a long way to go before we can truly say that engines can
be accurately designed. A lot of development work needs to be done in
order to produce a successful engine. Most of it has to be performed by
"greasy-fingered" engineers, who are often forced to work with trial-anderror
methods. They need all the useful help they can get from the various
specialties for reducing weight, improving fuel consumption, and increasing
both reliability and component life.
With respect to the biography, observe first that research is continually
being directed toward improving the capabilities of aircraft and engines.
General analyses of these results are periodically made to examine the
feasibility of improving existing airplanes, of making an airplane that will
either perform a useful service more economically, or of providing one that
was hitherto unavailable.
When the feasibility of an improved engine/airframe system is determined,
an analysis similar to that presented in Chapter 3 is initiated.
Engines and airplanes having the proposed advanced concepts are simulated
on large digital computers. Estimates are made of the properties and
behavior of every significant part of the engine and aircraft. Vital features
of the synthesized airplane are calculated along its proposed flight paths.
Numerous modifications of the synthesized engines and aircraft are examined
until the calculated optimum reliable airframe/engine combination for
the proposed mission is found.
The results are then reviewed to determine the worthiness of the airplane.
This review includes an estimate of the cost of designing and developing the
engine and airplane, the cost of manufacturing the desired number, and the
cost of operating and maintaining them over their expected life. These
economic evaluations (which are often identified by the words "cost of
ownership," "life-cycle-costs," or "return on investment") are repeated
throughout the life of the enterprise.
If a decision is made to proceed with the design and development of an
engine, the initial specifications are provided by the preceding studies. Tests
are begun on the various components to evaluate their effectiveness: i.e.,
how does the performance compare to that expected? Several design
modifications are frequently required to realize the initial expectations.
After the components are functioning reasonably well, they are assembled
into an engine to evaluate the overall performance and reliability.
Parts of new engines usually operate in more hostile environments than any
of their predecessors and unanticipated interactions among the parts (both
aerothermodynamic and mechanical) are occasionally encountered. Many
events can and do happen. Some problems can be rectified by minor
changes in the engine design. Others may require major redesigns that need
to be re-evaluated on component rigs. In some instances, added required
costs of time and money can put the project in jeopardy.
When the engine performs its functions on a test stand with sufficient
reliability, it is ready for flight tests. New interactions between the airframe
and the engine introduce additional problems. This is particularly true
when there is no previous experience with the maneuvers and flight regimes
demanded of the airplane. Expedited rebirth and development of critical
engine components have been necessary even at this late phase of development.
Dedicated and competent engineering teamwork usually solves the
problems, finally producing an airplane that meets its specifications.
After an engine has been certified by a government bureau, it enters its
intended service. When many hours of flight have elapsed, new major
problems have suddenly appeared. Although most of these are mechanical
(related to fatigue), they probably represent a poor tradeoff, initially made
with incomplete data, between aerodynamics and stress. New evaluations
and new designs have to be made and developed in a hurry to solve the
problems. Again, it is the teamwork of many talents that overcome the
difficulties.
This concluding section of the chapter is presented to show that all
specialties have a long way to go before we can truly say that engines can
be accurately designed. A lot of development work needs to be done in
order to produce a successful engine. Most of it has to be performed by
"greasy-fingered" engineers, who are often forced to work with trial-anderror
methods. They need all the useful help they can get from the various
specialties for reducing weight, improving fuel consumption, and increasing
both reliability and component life.
With respect to the biography, observe first that research is continually
being directed toward improving the capabilities of aircraft and engines.
General analyses of these results are periodically made to examine the
feasibility of improving existing airplanes, of making an airplane that will
either perform a useful service more economically, or of providing one that
was hitherto unavailable.
When the feasibility of an improved engine/airframe system is determined,
an analysis similar to that presented in Chapter 3 is initiated.
Engines and airplanes having the proposed advanced concepts are simulated
on large digital computers. Estimates are made of the properties and
behavior of every significant part of the engine and aircraft. Vital features
of the synthesized airplane are calculated along its proposed flight paths.
Numerous modifications of the synthesized engines and aircraft are examined
until the calculated optimum reliable airframe/engine combination for
the proposed mission is found.
The results are then reviewed to determine the worthiness of the airplane.
This review includes an estimate of the cost of designing and developing the
engine and airplane, the cost of manufacturing the desired number, and the
cost of operating and maintaining them over their expected life. These
economic evaluations (which are often identified by the words "cost of
ownership," "life-cycle-costs," or "return on investment") are repeated
throughout the life of the enterprise.
If a decision is made to proceed with the design and development of an
engine, the initial specifications are provided by the preceding studies. Tests
are begun on the various components to evaluate their effectiveness: i.e.,
how does the performance compare to that expected? Several design
modifications are frequently required to realize the initial expectations.
After the components are functioning reasonably well, they are assembled
into an engine to evaluate the overall performance and reliability.
Parts of new engines usually operate in more hostile environments than any
of their predecessors and unanticipated interactions among the parts (both
aerothermodynamic and mechanical) are occasionally encountered. Many
events can and do happen. Some problems can be rectified by minor
changes in the engine design. Others may require major redesigns that need
to be re-evaluated on component rigs. In some instances, added required
costs of time and money can put the project in jeopardy.
When the engine performs its functions on a test stand with sufficient
reliability, it is ready for flight tests. New interactions between the airframe
and the engine introduce additional problems. This is particularly true
when there is no previous experience with the maneuvers and flight regimes
demanded of the airplane. Expedited rebirth and development of critical
engine components have been necessary even at this late phase of development.
Dedicated and competent engineering teamwork usually solves the
problems, finally producing an airplane that meets its specifications.
After an engine has been certified by a government bureau, it enters its
intended service. When many hours of flight have elapsed, new major
problems have suddenly appeared. Although most of these are mechanical
(related to fatigue), they probably represent a poor tradeoff, initially made
with incomplete data, between aerodynamics and stress. New evaluations
and new designs have to be made and developed in a hurry to solve the
problems. Again, it is the teamwork of many talents that overcome the
difficulties.
AIRCRAFT ADVANCED FLOW CALCULATIONS
ADVANCED FLOW CALCULATIONS
The art of flow calculations is advancing at a rapid rate. The revolution
in digital computers has inspired an accompanying revolution in the art of
numerical analysis. Even a lengthy treatise at this point would do an
injustice to this subject. Just enough information will be given to describe
what lies behind some of the ongoing activities.
Most of the new procedures embody some form of the Navier-Stokes
equations that include the Reynolds turbulence stresses. Some models for
estimating the distribution of turbulence are also required. The available
models are approximate and subject to opinions and improvements. Nevertheless,
some very useful advances have been made with the crude turbulence
models now existing. Several methods of calculating a flowfield can be
developed from this base. Some procedures use the concept of boundary
layer.
Another technique that was recently introduced borrows an idea from
electromagnetic theory. The flow vector is divided into two components--
one with no curl and the other with no divergence. The latter vector
describes all the vorticity in the flow. In some respects, it is really a
generalization of the boundary-layer concept. Compatibility conditions
between the two vector fields are required. The first field is mixed elliptichyperbolic
and can be solved by known "relaxation" procedures. The
second field is essentially parabolic and is solved by "marching methods."
(An upstream vector does not know what a downstream vector is doing.)
Two classes of computational techniques are being pursued. One class
begins by assuming the flow to be at rest and imposes the boundary
conditions. The time-unsteady equations of motion then estimate the
acceleration of the flow until it reaches equilibrium. The other class uses
finite-difference or finite-element methods. Transformation of coordinates
as a function of the local Mach number and the location of the point of
interest in the flowfield may be used. Methods of taking the finite differences
or of describing the finite elements also vary with both the local Mach
number and the point in the field.
All the programs are very involved and their preparation presently
requires an outstanding ability in computer programming, numerical analysis,
and fluid mechanics. Because of the enormity of the problem, existing
programs must be considered to be only partially developed. Even in this
crude state, their use has, for example, indicated ways of designing turbine
vanes so that the secondary-flow losses are reduced. The experimental test
of the resulting turbine was more than gratifying.
This field is moving forward. Useful new concepts are continually being
disclosed. As a result, the ability to accurately analyze complex flows in
detail is noticeably improving from year to year. Close attention must be
paid to this activity.
The pursuit of advanced three-dimensional analysis will bring about step
improvements in turbomachinery performance. Side benefits will be the
reduction of expensive testing and the delineation of forcing functions that
affect blade vibration
The art of flow calculations is advancing at a rapid rate. The revolution
in digital computers has inspired an accompanying revolution in the art of
numerical analysis. Even a lengthy treatise at this point would do an
injustice to this subject. Just enough information will be given to describe
what lies behind some of the ongoing activities.
Most of the new procedures embody some form of the Navier-Stokes
equations that include the Reynolds turbulence stresses. Some models for
estimating the distribution of turbulence are also required. The available
models are approximate and subject to opinions and improvements. Nevertheless,
some very useful advances have been made with the crude turbulence
models now existing. Several methods of calculating a flowfield can be
developed from this base. Some procedures use the concept of boundary
layer.
Another technique that was recently introduced borrows an idea from
electromagnetic theory. The flow vector is divided into two components--
one with no curl and the other with no divergence. The latter vector
describes all the vorticity in the flow. In some respects, it is really a
generalization of the boundary-layer concept. Compatibility conditions
between the two vector fields are required. The first field is mixed elliptichyperbolic
and can be solved by known "relaxation" procedures. The
second field is essentially parabolic and is solved by "marching methods."
(An upstream vector does not know what a downstream vector is doing.)
Two classes of computational techniques are being pursued. One class
begins by assuming the flow to be at rest and imposes the boundary
conditions. The time-unsteady equations of motion then estimate the
acceleration of the flow until it reaches equilibrium. The other class uses
finite-difference or finite-element methods. Transformation of coordinates
as a function of the local Mach number and the location of the point of
interest in the flowfield may be used. Methods of taking the finite differences
or of describing the finite elements also vary with both the local Mach
number and the point in the field.
All the programs are very involved and their preparation presently
requires an outstanding ability in computer programming, numerical analysis,
and fluid mechanics. Because of the enormity of the problem, existing
programs must be considered to be only partially developed. Even in this
crude state, their use has, for example, indicated ways of designing turbine
vanes so that the secondary-flow losses are reduced. The experimental test
of the resulting turbine was more than gratifying.
This field is moving forward. Useful new concepts are continually being
disclosed. As a result, the ability to accurately analyze complex flows in
detail is noticeably improving from year to year. Close attention must be
paid to this activity.
The pursuit of advanced three-dimensional analysis will bring about step
improvements in turbomachinery performance. Side benefits will be the
reduction of expensive testing and the delineation of forcing functions that
affect blade vibration
INTERACTION WITH OTHER SPECIALTIES
INTERACTION WITH OTHER SPECIALTIES
It cannot be mentioned too often that there are many disciplines involved
in designing, developing, and manufacturing engines. How well one discipline
understands the problems and technology of other disciplines plays an
important role in determining costs as well as the elapsed time between the
initiation of design and the delivery of an approved product.
An aerodynamicist's lack of appreciation of stress and material technology
may require him to design and redesign a blade a number of times
before the aerodynamic and endurance requirements are satisfied. If manufacturing
technology is ignored, an unnecessarily expensive process may
have to be devised to make and inspect the blade. It is especially note
worthy that a part which can be thoroughly inspected is often preferred to
a potentially superior part which does not lend itself to inspection. Frequently,
the potentially superior part turns out to be inferior because it
does not conform to design requirements. Simple design changes that
improve the ability to reproduce a product, perhaps with a slight penalty in
latent performance, should always be in the thoughts of an aerodynamic
designer. This emphasizes the requirement for the designer to understand
the real needs of the customer.
As mentioned in Sec. 1.2, the purpose of an airplane is to render a service
that enough people want and can afford. The costs to the eventual
customer include his share of the expenses associated with the initial design
and development, manufacturing, maintenance, and availability. (Availability
signifies, among other connotations, whether nine or ten engines
must be purchased to be sure that at least eight are available when needed.)
By understanding the total picture, an honest evaluation can be made
about whether the added cost of a supposedly more efficient part is worth
its latent contribution to reduced fuel consumption, reduced weight, or
increased thrust. This knowledge, of course, cannot be acquired instantly.
The successful aerodynamic designer will always be alert to opportunities
that will improve his understanding of these interrelationships.
Besides adopting this long-range philosophy, there are many areas where
a good understanding of peripheral technology should be sought almost
immediately. A few subjects and problem areas have been selected to
illustrate the need for the aerodynamicist to be involved in many activities.
The important subject of controls is not covered here since some attention
was directed to them earlier.
Dimensional Integrity
Notice was taken in previous sections of some effects of the dimensional
changes that accompany temperature variations throughout an engine. The
existence of variations in temperature and dimensional changes along the
length of the engine during steady-state operation are readily appreciated.
The effects of transient operation on changing these variations of temperature
and dimensions with time should also be recognized.
Circumferential variations in temperature near an outer casing can cause
bulges in the casing, with a resulting local increase in clearance. Simple
calculations can indicate how easily noticeable increases in such clearances
are achieved.
It is practically impossible to keep uniform clearance between stationary
and rotating parts at all times. There is, however, always some operating
condition and one or more areas of an engine where the clearance is a
minimum. The circumstances depend upon the temperatures and vibrations.
Airplane maneuvers are also involved. The requirement of mechanical
integrity sets the magnitude of the minimum clearances. The clearances
at other operating conditions are then automatically defined.
The relative growth of the rotating and stationary parts depends to some
extent on the aerothermodynamic design. Heat-transfer rates and blade
shapes are involved.
It cannot be mentioned too often that there are many disciplines involved
in designing, developing, and manufacturing engines. How well one discipline
understands the problems and technology of other disciplines plays an
important role in determining costs as well as the elapsed time between the
initiation of design and the delivery of an approved product.
An aerodynamicist's lack of appreciation of stress and material technology
may require him to design and redesign a blade a number of times
before the aerodynamic and endurance requirements are satisfied. If manufacturing
technology is ignored, an unnecessarily expensive process may
have to be devised to make and inspect the blade. It is especially note
worthy that a part which can be thoroughly inspected is often preferred to
a potentially superior part which does not lend itself to inspection. Frequently,
the potentially superior part turns out to be inferior because it
does not conform to design requirements. Simple design changes that
improve the ability to reproduce a product, perhaps with a slight penalty in
latent performance, should always be in the thoughts of an aerodynamic
designer. This emphasizes the requirement for the designer to understand
the real needs of the customer.
As mentioned in Sec. 1.2, the purpose of an airplane is to render a service
that enough people want and can afford. The costs to the eventual
customer include his share of the expenses associated with the initial design
and development, manufacturing, maintenance, and availability. (Availability
signifies, among other connotations, whether nine or ten engines
must be purchased to be sure that at least eight are available when needed.)
By understanding the total picture, an honest evaluation can be made
about whether the added cost of a supposedly more efficient part is worth
its latent contribution to reduced fuel consumption, reduced weight, or
increased thrust. This knowledge, of course, cannot be acquired instantly.
The successful aerodynamic designer will always be alert to opportunities
that will improve his understanding of these interrelationships.
Besides adopting this long-range philosophy, there are many areas where
a good understanding of peripheral technology should be sought almost
immediately. A few subjects and problem areas have been selected to
illustrate the need for the aerodynamicist to be involved in many activities.
The important subject of controls is not covered here since some attention
was directed to them earlier.
Dimensional Integrity
Notice was taken in previous sections of some effects of the dimensional
changes that accompany temperature variations throughout an engine. The
existence of variations in temperature and dimensional changes along the
length of the engine during steady-state operation are readily appreciated.
The effects of transient operation on changing these variations of temperature
and dimensions with time should also be recognized.
Circumferential variations in temperature near an outer casing can cause
bulges in the casing, with a resulting local increase in clearance. Simple
calculations can indicate how easily noticeable increases in such clearances
are achieved.
It is practically impossible to keep uniform clearance between stationary
and rotating parts at all times. There is, however, always some operating
condition and one or more areas of an engine where the clearance is a
minimum. The circumstances depend upon the temperatures and vibrations.
Airplane maneuvers are also involved. The requirement of mechanical
integrity sets the magnitude of the minimum clearances. The clearances
at other operating conditions are then automatically defined.
The relative growth of the rotating and stationary parts depends to some
extent on the aerothermodynamic design. Heat-transfer rates and blade
shapes are involved.
INTERRELATIONS AMONG AERODYNAMIC COMPONENTS
INTERRELATIONS AMONG AERODYNAMIC COMPONENTS
The other chapters of these volumes describe how the design of engine
components may be implemented once the engine type is decided and the
pressure ratios and turbine inlet temperatures have been tentatively selected.
Further elaboration is not required here. Some observations of
mutual interactions among the components remain to be discussed, as well
as some incidental information about individual components that influences
their effectiveness.
Most of the problems in this category were anticipated in the discussions
of the previous sections. Any component that delivers a nonuniform
distribution to components downstream contributes losses and other problems
to those components. Conversely, events in a downstream component
may affect an upstream one. Because of these interactions, potentially
beneficial concepts for one component sometimes have to be deferred to
help another.
Several of the authors in this series have emphasized the close cooperation
among the many disciplines that is mandatory if a design is to be
efficiently executed. Pertinent topics have been discussed in Refs. 1 and 2,
and in the remaining chapters of this volume. One object of this section is
to identify some of the compromises involved in selecting turbine inlet
temperatures, rotor speeds, and rotor diameters. Important interactions
among the aerodynamic components are also reviewed in this section.
Temperature and Speed
Most modern engines can profitably use the highest achievable turbine
inlet temperatures. An obvious exception is a jet engine for subsonic or
low-supersonic speeds. If the bypass ratio of the fans is limited by problems
such as ground clearance, Eq. (1.20) shows that aerothermodynamics can
again constrain the desired temperature. In the other cases, temperature is
limited by material strength, thermal corrosion, and aerodynamic and
cooling losses necessary for surviving high temperatures.
Section 4.3 of Ref. 2 offers a good discussion of the design restraints
imposed by temperature and stress. The same chapter shows how stress
problems are eased by cooling; it also shows how the act of cooling
introduces new losses and how the blade shapes required for cooling invite
still further losses. Chapter 5 of the same reference provides further
material on this subject; Sec. 5.2 in particular gives a good description of
the problems of designing for high temperature.
An effect of temperature on material strength is quantified by Fig. 1.38.
The abscissa is the Larsen-Miller parameter, which correlates the combined
effects of the level and duration of stress and temperature. The ordinate
shows the stress allowed for each value of the abscissa.
Two properties are noteworthy. Temperature has an almost exponential
effect on useful life--the life is about halved for a 25°F increase in
temperature. The second feature concerns the uncertainty of the data. The
discussions cited in Ref. 2 provide evidence about the lack of uniformity in
the temperatures approaching turbines. An educated guess is necessary to
avoid either an unreliable or uncompetitive design. The indicated uncertainty
of the material data is another challenge. Some designers use the
three-sigma data and add an arbitrary value, e.g., 50°F, to the expected
temperature. Since life is so sensitive to temperature, experience with
similar tasks and with one's associates is a valuable asset for making proper
choices.
Figure 1.38 alone would strongly support a choice of low rotor speeds to
avoid high stresses. Low linear and angular blade speeds would result from
this decision. Euler's turbine equation, however, leads to the observation
that power is proportional to the cube of the linear blade speeds and that
high speeds are required for lightweight units. These conflicts are resolved
by making many design iterations to arrive at an attractive balance.
Experience with other engines is again an asset.
Another restraint on speeds is the Mach number relative to the rotors of
fan and compressor blades. This problem is discussed in Chapter 3 of Ref.
2. Fan blade stresses are also a determinant of allowable speed, particularly
when the added weight of part span shrouds (or dampers or clappers,
visible in Fig. 1.27) are used as protection from flutter or to re-enforce the
blades in the event of an impact with a large bird during flight.
Diameters
The diameters of the components are determined principally by the gas
flow rate. Inasmuch as the radial force on a blade is proportional to the
flow area, there is a natural desire to keep this area as low as possible by
maintaining the axial component of the Mach numbers at high subsonic
values.
This solution opens the door to other problems, however. When a
one-dimensional calculation indicates that the axial component of the
Mach number leaving a blade row is about 0.6, the row is probably at
limiting loading. What happens is that the effective axial component of the
Mach number is about 1--the wakes from the blades and the combined
thicknesses of the boundary layers at the tip and casing contract the
effective flow area by the 19% needed to account for the difference.
The Mach number of the flow at the compressor inlet affects the relative
position of the corrected speed lines; see, e.g., Fig. 1.12a. The gas speed is
more or less proportional to the rotating speed, but the rate of change of (WvfT/P) with the airspeed decreases to zero as the throughflow Mach
number approaches 1. (See Fig. 1.39.) The rate of change of corrected
weight flow behaves the same way. This accounts for the crowding of the
speed lines in Fig. 1.12a at their higher values. This situation can be
troublesome and must be avoided.
There are other reasons for using less than the maximum flow rates.
Efficiency suffers when the blade heights become too small and the area of
the flow boundaries offering nothing but friction becomes comparable to or
greater than those areas responsible for the transfer of useful energy.
Again, high velocities approaching combustors must be diffused at the
expense of space and, sometimes, available energy. Moreover, a large
acceleration of flow ahead of an engine nozzle throat usually aids the thrust
coefficient by improving the flow distribution; low upstream velocities are
then beneficial. These are a few of the reasons why diameters are also
selected only after an extensive inquiry of the design iterations.
The other chapters of these volumes describe how the design of engine
components may be implemented once the engine type is decided and the
pressure ratios and turbine inlet temperatures have been tentatively selected.
Further elaboration is not required here. Some observations of
mutual interactions among the components remain to be discussed, as well
as some incidental information about individual components that influences
their effectiveness.
Most of the problems in this category were anticipated in the discussions
of the previous sections. Any component that delivers a nonuniform
distribution to components downstream contributes losses and other problems
to those components. Conversely, events in a downstream component
may affect an upstream one. Because of these interactions, potentially
beneficial concepts for one component sometimes have to be deferred to
help another.
Several of the authors in this series have emphasized the close cooperation
among the many disciplines that is mandatory if a design is to be
efficiently executed. Pertinent topics have been discussed in Refs. 1 and 2,
and in the remaining chapters of this volume. One object of this section is
to identify some of the compromises involved in selecting turbine inlet
temperatures, rotor speeds, and rotor diameters. Important interactions
among the aerodynamic components are also reviewed in this section.
Temperature and Speed
Most modern engines can profitably use the highest achievable turbine
inlet temperatures. An obvious exception is a jet engine for subsonic or
low-supersonic speeds. If the bypass ratio of the fans is limited by problems
such as ground clearance, Eq. (1.20) shows that aerothermodynamics can
again constrain the desired temperature. In the other cases, temperature is
limited by material strength, thermal corrosion, and aerodynamic and
cooling losses necessary for surviving high temperatures.
Section 4.3 of Ref. 2 offers a good discussion of the design restraints
imposed by temperature and stress. The same chapter shows how stress
problems are eased by cooling; it also shows how the act of cooling
introduces new losses and how the blade shapes required for cooling invite
still further losses. Chapter 5 of the same reference provides further
material on this subject; Sec. 5.2 in particular gives a good description of
the problems of designing for high temperature.
An effect of temperature on material strength is quantified by Fig. 1.38.
The abscissa is the Larsen-Miller parameter, which correlates the combined
effects of the level and duration of stress and temperature. The ordinate
shows the stress allowed for each value of the abscissa.
Two properties are noteworthy. Temperature has an almost exponential
effect on useful life--the life is about halved for a 25°F increase in
temperature. The second feature concerns the uncertainty of the data. The
discussions cited in Ref. 2 provide evidence about the lack of uniformity in
the temperatures approaching turbines. An educated guess is necessary to
avoid either an unreliable or uncompetitive design. The indicated uncertainty
of the material data is another challenge. Some designers use the
three-sigma data and add an arbitrary value, e.g., 50°F, to the expected
temperature. Since life is so sensitive to temperature, experience with
similar tasks and with one's associates is a valuable asset for making proper
choices.
Figure 1.38 alone would strongly support a choice of low rotor speeds to
avoid high stresses. Low linear and angular blade speeds would result from
this decision. Euler's turbine equation, however, leads to the observation
that power is proportional to the cube of the linear blade speeds and that
high speeds are required for lightweight units. These conflicts are resolved
by making many design iterations to arrive at an attractive balance.
Experience with other engines is again an asset.
Another restraint on speeds is the Mach number relative to the rotors of
fan and compressor blades. This problem is discussed in Chapter 3 of Ref.
2. Fan blade stresses are also a determinant of allowable speed, particularly
when the added weight of part span shrouds (or dampers or clappers,
visible in Fig. 1.27) are used as protection from flutter or to re-enforce the
blades in the event of an impact with a large bird during flight.
Diameters
The diameters of the components are determined principally by the gas
flow rate. Inasmuch as the radial force on a blade is proportional to the
flow area, there is a natural desire to keep this area as low as possible by
maintaining the axial component of the Mach numbers at high subsonic
values.
This solution opens the door to other problems, however. When a
one-dimensional calculation indicates that the axial component of the
Mach number leaving a blade row is about 0.6, the row is probably at
limiting loading. What happens is that the effective axial component of the
Mach number is about 1--the wakes from the blades and the combined
thicknesses of the boundary layers at the tip and casing contract the
effective flow area by the 19% needed to account for the difference.
The Mach number of the flow at the compressor inlet affects the relative
position of the corrected speed lines; see, e.g., Fig. 1.12a. The gas speed is
more or less proportional to the rotating speed, but the rate of change of (WvfT/P) with the airspeed decreases to zero as the throughflow Mach
number approaches 1. (See Fig. 1.39.) The rate of change of corrected
weight flow behaves the same way. This accounts for the crowding of the
speed lines in Fig. 1.12a at their higher values. This situation can be
troublesome and must be avoided.
There are other reasons for using less than the maximum flow rates.
Efficiency suffers when the blade heights become too small and the area of
the flow boundaries offering nothing but friction becomes comparable to or
greater than those areas responsible for the transfer of useful energy.
Again, high velocities approaching combustors must be diffused at the
expense of space and, sometimes, available energy. Moreover, a large
acceleration of flow ahead of an engine nozzle throat usually aids the thrust
coefficient by improving the flow distribution; low upstream velocities are
then beneficial. These are a few of the reasons why diameters are also
selected only after an extensive inquiry of the design iterations.
LOSSES IN AVAILABLE ENERGY
LOSSES IN AVAILABLE ENERGY
The term losses describes the difference between the energy ideally
needed to propel an airplane a given distance and the latent energy of the
fuel actually consumed. The use of efficiencies has implied the existence of
losses without identifying their cause. Since minimizing losses is a major
function of engine design, it is now appropriate to review the sources of
various losses in engines. This is the purpose of this section.
Note first that the thermodynamic cycle used in propulsion engines has
an inherent loss mechanism that is not included under aerodynamics--the
fuel energy rendered unavailable because adiabatic combustion is necessary.
In comparison to a reversible isothermal process, this type of combustion
wastes about half of the potential energy of the fuel. This loss is not
recognized in Eq. (1.3).
The losses implied by the "ideal cycle efficiency," the quantity (1 - to/T3)
in Eq. (1.3), are not aerodynamic either. Neither are they included as part
of the compression or expansion efficiencies. It is the losses causing the
component efficiencies to be less than unity that are the subject of the rest
of this section, which first examines the requirements for a consistent
definition of losses and then discusses their sources.
The term losses describes the difference between the energy ideally
needed to propel an airplane a given distance and the latent energy of the
fuel actually consumed. The use of efficiencies has implied the existence of
losses without identifying their cause. Since minimizing losses is a major
function of engine design, it is now appropriate to review the sources of
various losses in engines. This is the purpose of this section.
Note first that the thermodynamic cycle used in propulsion engines has
an inherent loss mechanism that is not included under aerodynamics--the
fuel energy rendered unavailable because adiabatic combustion is necessary.
In comparison to a reversible isothermal process, this type of combustion
wastes about half of the potential energy of the fuel. This loss is not
recognized in Eq. (1.3).
The losses implied by the "ideal cycle efficiency," the quantity (1 - to/T3)
in Eq. (1.3), are not aerodynamic either. Neither are they included as part
of the compression or expansion efficiencies. It is the losses causing the
component efficiencies to be less than unity that are the subject of the rest
of this section, which first examines the requirements for a consistent
definition of losses and then discusses their sources.
Multishaft Turboprop Engine
Multishaft Turboprop Engine
It is obvious that the multishaft concepts can be extended to a turboprop
engine with the propeller turbine replacing the role of the fan and most of
the role of the thrust nozzle. The off-design behavior would be similar to
that discussed in the previous section on the two-shaft jet engine.
Afterburning
A photograph of a single-shaft turbojet with afterburning was presented
in Fig. 1.15. Two designs of two-shaft turbofan engines with afterburners
are shown in Fig. 1.29. Notice that the gas stream from the engine is mixed
with that of the fan ahead of the afterburner. The static pressures of both
streams must be the same where they merge, e.g., at point A in the upper
figure. This requirement affects the off-design performance of both the fan
and gas generator, but the general principles previously noted still apply.
Multistage fans and afterburners typify configurations needed by supersonic
aircraft. Except for their proportions, the rotating structures resemble
those of Fig. 1.27. The function of the mixers just ahead of the afterburners
is explained in Chapter 2 of Ref. 2.
Igniting an afterburner poses a special off-design and control problem
that is treated in Chapter 2 of Ref. 2. In brief, the effective Mach number
of the flow in the nozzle throat is unity during normal operation. When the
gas temperature is suddenly raised, the nozzle throat area must rapidly
increase to pass the mass flow without an excessive pressure rise. The rate
of change of the fuel flow must be accurately synchronized with the rate of
change of the throat area. If the area increases too rapidly, thrust is
momentarily lost just when an increase is demanded. If it is opened too
slowly, the fan and possibly the compressor are pushed into surge. The
magnitude of the control problem varies inversely with the permissible
surge margins of the compressors.
It is obvious that the multishaft concepts can be extended to a turboprop
engine with the propeller turbine replacing the role of the fan and most of
the role of the thrust nozzle. The off-design behavior would be similar to
that discussed in the previous section on the two-shaft jet engine.
Afterburning
A photograph of a single-shaft turbojet with afterburning was presented
in Fig. 1.15. Two designs of two-shaft turbofan engines with afterburners
are shown in Fig. 1.29. Notice that the gas stream from the engine is mixed
with that of the fan ahead of the afterburner. The static pressures of both
streams must be the same where they merge, e.g., at point A in the upper
figure. This requirement affects the off-design performance of both the fan
and gas generator, but the general principles previously noted still apply.
Multistage fans and afterburners typify configurations needed by supersonic
aircraft. Except for their proportions, the rotating structures resemble
those of Fig. 1.27. The function of the mixers just ahead of the afterburners
is explained in Chapter 2 of Ref. 2.
Igniting an afterburner poses a special off-design and control problem
that is treated in Chapter 2 of Ref. 2. In brief, the effective Mach number
of the flow in the nozzle throat is unity during normal operation. When the
gas temperature is suddenly raised, the nozzle throat area must rapidly
increase to pass the mass flow without an excessive pressure rise. The rate
of change of the fuel flow must be accurately synchronized with the rate of
change of the throat area. If the area increases too rapidly, thrust is
momentarily lost just when an increase is demanded. If it is opened too
slowly, the fan and possibly the compressor are pushed into surge. The
magnitude of the control problem varies inversely with the permissible
surge margins of the compressors.
Multishaft Turbofan Engines
Multishaft Turbofan Engines
An arrangement of this type of engine is suggested by the sketch of Fig.
1.26. A photograph of a production version for subsonic flight is shown in
Fig. 1.27. Recall that the fan is nothing more than a compressor producing
a relatively low pressure ratio. The off-design behavior of the components
of this engine is similar to the two-shaft turbojet, since the turbines and jet
divide the pressure ratio in about the same way. Observe that when the
ETR is lowered, the power to the jet diminishes faster than that to the fan.
The discussion of turbofan engines in Sec. 1.3 demonstrated that there is an
optimum division of these powers, which depends on flight speed and
internal efficiencies. Prolonged operation at two or more flight conditions
requires the power division to be compromised unless some variable feature
is provided.
Another engine currently in production uses three concentric shafts: there
are two sets of compressor and turbine rotors and a turbine on the third
shaft drives the fan. The thrust nozzle is downstream of the third turbine.
Reducing the ETR first causes the nozzle to unchoke; the fan turbine
becomes unchoked at a slightly lower ETR. The changes in the ratio of fan
power to jet power is different for this design and a different off-design
efficiency is thus expected.
Still another concept is shown in Fig. 1.28. The objective of the design is
to put the turbine driving the fan immediately downstream of the highpressure
turbine; the tortuous flow path was accepted in order to achieve
this objective. This arrangement was a deliberate attempt to provide a
favorable balance of fan and jet power at low levels of specific engine
power, while still maintaining acceptable engine efficiencies. The aerodynamic
aims were almost realized in spite of the intricate flow path. There
was an unfortunate mechanical problem, however. The regions of high
pressure within the engine are on the right side of the photograph for both
the compressors and turbines. A large pneumatic force, which could not be
accommodated by rolling element thrust bearings, is exerted toward the left
on the rotors. Pressurized air bearings requiring rotating air seals having
large diameters are necessary. Precise control of clearances is mandatory to
prevent the associated leakage losses from overwhelming the gains; thus,
the costs of manufacturing and maintenance control the value of this
design. This is a good illustration of a need for close cooperation among
aerodynamic, mechanical design, and manufacturing specialists early in a
program.
An arrangement of this type of engine is suggested by the sketch of Fig.
1.26. A photograph of a production version for subsonic flight is shown in
Fig. 1.27. Recall that the fan is nothing more than a compressor producing
a relatively low pressure ratio. The off-design behavior of the components
of this engine is similar to the two-shaft turbojet, since the turbines and jet
divide the pressure ratio in about the same way. Observe that when the
ETR is lowered, the power to the jet diminishes faster than that to the fan.
The discussion of turbofan engines in Sec. 1.3 demonstrated that there is an
optimum division of these powers, which depends on flight speed and
internal efficiencies. Prolonged operation at two or more flight conditions
requires the power division to be compromised unless some variable feature
is provided.
Another engine currently in production uses three concentric shafts: there
are two sets of compressor and turbine rotors and a turbine on the third
shaft drives the fan. The thrust nozzle is downstream of the third turbine.
Reducing the ETR first causes the nozzle to unchoke; the fan turbine
becomes unchoked at a slightly lower ETR. The changes in the ratio of fan
power to jet power is different for this design and a different off-design
efficiency is thus expected.
Still another concept is shown in Fig. 1.28. The objective of the design is
to put the turbine driving the fan immediately downstream of the highpressure
turbine; the tortuous flow path was accepted in order to achieve
this objective. This arrangement was a deliberate attempt to provide a
favorable balance of fan and jet power at low levels of specific engine
power, while still maintaining acceptable engine efficiencies. The aerodynamic
aims were almost realized in spite of the intricate flow path. There
was an unfortunate mechanical problem, however. The regions of high
pressure within the engine are on the right side of the photograph for both
the compressors and turbines. A large pneumatic force, which could not be
accommodated by rolling element thrust bearings, is exerted toward the left
on the rotors. Pressurized air bearings requiring rotating air seals having
large diameters are necessary. Precise control of clearances is mandatory to
prevent the associated leakage losses from overwhelming the gains; thus,
the costs of manufacturing and maintenance control the value of this
design. This is a good illustration of a need for close cooperation among
aerodynamic, mechanical design, and manufacturing specialists early in a
program.
Two-Shaft Jet Engine
Two-Shaft Jet Engine
Simplified analysis. The maximum practical pressure, or temperature,
ratio for the compressor of a single-shaft jet engine is limited. The volume
flow rate at the inlet determines the diameter and the rotating speed is
limited by the highest allowable Much number at the inlet stage. As the
pressure increases within a compressor, the volume flow rate reduces and
the air temperature rises. The hydraulic diameters of the flow passages
continuously decrease from inlet to outlet and the corresponding pressure
ratio produced by each stage becomes smaller and smaller. Eventually, the
unit becomes heavy, inefficient, and impractical.
The increase in air temperature through a compressor lowers the Much
numbers of the flow about the blades. It is thus reasonable to use two
compressors; then, after a moderate pressure ratio is developed, the flow is
passed to another compressor having a smaller diameter and a greater
rotating speed. In the same way, the use of two turbines also offers
advantages. Designs with two separate shafts, as shown in Fig. 1.22, are
thus recommended. The reduced size of the inner spool offers the additional
benefit of a reduction in weight. Although the pressure ratio of each
compressor is smaller than that used in Fig. 1.12a, the overall pressure ratio
and efficiency can be increased markedly. Figure 1.23 presents typical maps
of the compressors. The trends of the turbine maps of Fig. 1.13 are still
appropriate.
The hierarchy of pressure ratios across the turbine is similar to that of
the gas generators previously described. An example is shown in Fig. 1.24
for several overall pressure ratios up to 8 across the turbines and engine
nozzle. A logarithmic scale is used for the ordinate so that the constantpressure
ratios across components appear as parallel lines. All elements
were assumed to be choked at the arbitrary overall pressure ratio of 8. The
low-pressure turbine begins to unchoke when the pressure ratio is 6, while
the high-pressure turbine is practically choked down to an overall pressure
ratio of 3.4.
Simplified analysis. The maximum practical pressure, or temperature,
ratio for the compressor of a single-shaft jet engine is limited. The volume
flow rate at the inlet determines the diameter and the rotating speed is
limited by the highest allowable Much number at the inlet stage. As the
pressure increases within a compressor, the volume flow rate reduces and
the air temperature rises. The hydraulic diameters of the flow passages
continuously decrease from inlet to outlet and the corresponding pressure
ratio produced by each stage becomes smaller and smaller. Eventually, the
unit becomes heavy, inefficient, and impractical.
The increase in air temperature through a compressor lowers the Much
numbers of the flow about the blades. It is thus reasonable to use two
compressors; then, after a moderate pressure ratio is developed, the flow is
passed to another compressor having a smaller diameter and a greater
rotating speed. In the same way, the use of two turbines also offers
advantages. Designs with two separate shafts, as shown in Fig. 1.22, are
thus recommended. The reduced size of the inner spool offers the additional
benefit of a reduction in weight. Although the pressure ratio of each
compressor is smaller than that used in Fig. 1.12a, the overall pressure ratio
and efficiency can be increased markedly. Figure 1.23 presents typical maps
of the compressors. The trends of the turbine maps of Fig. 1.13 are still
appropriate.
The hierarchy of pressure ratios across the turbine is similar to that of
the gas generators previously described. An example is shown in Fig. 1.24
for several overall pressure ratios up to 8 across the turbines and engine
nozzle. A logarithmic scale is used for the ordinate so that the constantpressure
ratios across components appear as parallel lines. All elements
were assumed to be choked at the arbitrary overall pressure ratio of 8. The
low-pressure turbine begins to unchoke when the pressure ratio is 6, while
the high-pressure turbine is practically choked down to an overall pressure
ratio of 3.4.
Free Turbine Turboprop
Free Turbine Turboprop
Figure 1.21 differs from Fig. 1.20 in that the turbine functions are
separated--one turbine drives a compressor and another the propeller. The
turbine that drives the compressor behaves as it does in a single-shaft
turbojet and the propeller turbine has the pressure ratio and flow characteristics
of a nozzle. Curve C-C of Fig. 1.16 again represents an operating line.
This engine offers better efficiency at reduced specific power than the
single-shaft engine running at constant speed. The speed of the gas generator
is reduced when the ETR is lowered; however, since there is no
propeller or large gear train on this shaft, the inertia is comparatively low.
The finite time required to respond to a demand for more power is small
and usually not objectionable.
The free turbine engine is the preferred choice for helicopters, which
require the main rotor to spin freely if the engine should stop. The load
imposed by the turbine, gears, and accessories is small enough to allow this
to happen without the use of a clutch or a free-wheeling device.
When the engine is idling, both the engine and propeller speeds are low.
This is a desirable quality for ground operations because of the low noise
level. Low engine speeds at part power are also useful when extended
periods of time for loitering are necessary.
Figure 1.21 differs from Fig. 1.20 in that the turbine functions are
separated--one turbine drives a compressor and another the propeller. The
turbine that drives the compressor behaves as it does in a single-shaft
turbojet and the propeller turbine has the pressure ratio and flow characteristics
of a nozzle. Curve C-C of Fig. 1.16 again represents an operating line.
This engine offers better efficiency at reduced specific power than the
single-shaft engine running at constant speed. The speed of the gas generator
is reduced when the ETR is lowered; however, since there is no
propeller or large gear train on this shaft, the inertia is comparatively low.
The finite time required to respond to a demand for more power is small
and usually not objectionable.
The free turbine engine is the preferred choice for helicopters, which
require the main rotor to spin freely if the engine should stop. The load
imposed by the turbine, gears, and accessories is small enough to allow this
to happen without the use of a clutch or a free-wheeling device.
When the engine is idling, both the engine and propeller speeds are low.
This is a desirable quality for ground operations because of the low noise
level. Low engine speeds at part power are also useful when extended
periods of time for loitering are necessary.
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