THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS
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`M Rights Reserved (cid:9)
`
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`
`APPLICATION OF TRANSIENT AND DYNAMIC SIMULATIONS TO THE U.S. ARMY T55-L-712
`HELICOPTER ENGINE
`
`S. A. Savelle
`G.D. Garrard
`Sverdrup Technology, Inc. / AEDC Group
`1099 Avenue C
`Arnold AFB, TN 37389-9013
`
`111111111111a1111111111
`
`ABSTRACT
`The T55-L-712 turboshaft engine, used in the U.S. Army CH-
`47D Chinook helicopter, has been simulated using version 3.0 of the
`Advanced Turbine Engine Simulation Technique (ATEST) and ver-
`sion 1.0 of the Aerodynamic Turbine Engine Code (ATEC). The
`models simulate transient and dynamic engine operation from idle to
`maximum power and run on an IBM-compatible personal computer.
`ATEST is a modular one-dimensional component-level transient tur-
`bine engine simulation. The simulation is tailored to a specific engine
`using engine-specific component maps and an engine-specific super-
`visory subroutine that defines component interrelationships. A7EC is
`a one-dimensional, time-dependent, dynamic turbine engine simula-
`tion. ATEC simulates the operation of a gas turbine by solving the
`one-dimensional, time dependent Euler equations with turbomachin-
`ery source terms. The simulation uses elemental control volumes at
`the sub-component level (e.g. compressor stage).
`The paper discusses how limited information from a variety of
`sources was adapted for use in the T55 simulations and how common-
`ality between the models allowed reuse of the same material. The first
`application of a new turbine engine model, ATEC, to a specific engine
`is also discussed. Calibration and operational verification of the
`. simulations will be discussed, along with the status of the simulations.
`
`NOMEMCLATURE
`Specific Heat at Constant Pressure
`Cp (cid:9)
`Enthalpy
`Shaft Inertia
`Pressure
`Pressure Ratio
`Temperature
`Time
`Temperature Ratio
`Gas Row
`Airflow
`Engine Fuel Row
`
`WA
`WEE
`
`PR
`
`TR
`
`XN
`a
`
`8
`
`Shaft Speed
`Shaft Rotational Acceleration
`Ratio of Specific Heats
`Ratio of Pressure to Standard Pressure
`Ratio of Temperature to Standard Temperature
`Shaft Rotational Speed
`
`Subscript
`2 (cid:9)
`Compressor Inlet
`3 (cid:9)
`Compressor Exit
`4 (cid:9)
`Gas Generator Turbine Inlet
`Gas Generator Turbine Exit
`5
`Corrected to Standard Conditions
`Component Exit
`Gas Generator Shaft
`Component Inlet
`Power Turbine Shaft
`Reference for Normalization
`Standard Day Condition
`Total
`
`PT
`ref
`std
`
`PG
`
`INTRODUCTION
`At the Arnold Engineering Development Center (AEDC), as at
`other sea-level and altitude ground test facilities, turbine engines are
`routinely tested for operability characteristics throughout an engine's
`life cycle. Operability testing of an engine is performed during engine
`development and qualification to define engine stability limits, during
`production in response to problems in service, and during engine im-
`provement programs as part of requalification. • Test programs are
`performed at AEDC by turbine engine manufacturers and operators in
`support of all of these efforts. Increasingly high costs of engine test-
`ing and competitive pressures to reduce engineering turnaround times
`require that these test programs be designed to minimize time and
`costs while completing the program's objectives.
`
`The work reported herein was performed by the Arnold Engineering Development Center (AEDC), Air Force Materiel Command. Work and analysis for this research
`were done by personnel of Sverdrup Technology. Inc.. AEDC Group, technical services contractor. Further reproduction is authorized to satisfy the needs of the U. S.
`Government.
`
`Presented at the International Gas Turbine and Aemengine Congress & Exhibition
`Birmingham, UK — June 10-13, 1996
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`Computer simulations of the turbine engine are used in support
`of operability testing for test optimization and for test analysis. Tran-
`sient turbine engine models are used pretest to select required test
`conditions, as well as online and posttest for test data reduction and
`analysis. Fully dynamic turbine engine models are used as an analysis
`tool in the investigation of highly dynamic events such as surge and
`rotating stall and their effects on full engine operability. Dynamic
`models allow extrapolation of test data to areas of operation where
`testing of an actual engine is dangerous or destructive.
`The description of an engine model as transient or dynamic is
`subject to interpretation, so the two terms will be defined within the
`context of this paper to avoid confusion. A transient engine model
`solves the equations for conservation of mass and energy, and non-
`steady terms in the equation for conservation of momentum are ne-
`glected. Time-dependent compressibility effects within engine com-
`ponents are typically not modeled, so a transient cycle model can
`simulate unsteady engine behavior of approximately 20 Hz or less.
`The lumped-component cycle deck that simulates normal engine
`throttle and flight transients is typical of this type of model. A dy-
`namic engine model includes the equation for conservation of mo-
`mentum, and simulates unsteady engine behavior at frequencies up to
`approximately 300 Hz. This includes aerodynamic phenomena at a
`subcomponent level, such as inlet distortion, combustor instabilities,
`or compressor rotating stall or surge, as well as normal transient en-
`gine behavior.
`This paper describes the development of two T55-L-712 engine
`simulations for use in support of engine operability testing. The
`simulations use transient and dynamic mathematical engine models
`developed at AEDC and engine information from a variety of sources.
`The transient engine model, the Advanced Turbine Engine Simulation
`Technique (ATEST), is a mature model that has been in use since
`1991. The dynamic engine model, the Aerodynamic Turbine Engine
`Code (ATEC), has recently been developed at AEDC and the T55
`simulation is the first application of this model. The simulations are
`being used to support an experimental investigation by the Army Ve-
`hicle Propulsion Directorate (VPD) at the NASA Lewis Research
`Center to investigate the starting characteristics of the T55-L-712
`turboshaft engine.
`
`T55-L-712 ENGINE
`The T55 engine is a turboshaft gas turbine for a helicopter, with
`seven axial stages and one centrifugal stage of compression connected
`to a two-stage axial turbine in the gas generator. A compressor air
`bleed between the sixth and seventh axial stages is actuated for surge
`prevention during starts, accelerations and decelerations. The com-
`bustor is a reverse-flow annular design with atomizing fuel nozzles.
`Power is supplied to the helicopter through a concentric shaft con-
`nected to a two-stage axial fire power turbine. A cutaway schematic
`of the T55-L-712 is shown in Fig. I.
`The engine is controlled by a Hamilton Standard JFC31-22 hy-
`dromechanical fuel control. Through the helicopter control systems,
`the pilot sets two engine control lever inputs. The lever inputs are
`arbitrary angle values with a defined relationship to one or more
`• physical parameters. The first lever angle input, Alphal, provides the
`engine control system with the desired maximum gas generator speed.
`The second lever angle input, Alpha2, provides the engine control
`system with the desired power turbine torque and speed. In addi-
`tion to pilot inputs, the engine control uses four engine feedback pa-
`rameters: XN1 and XN2, gas generator and power turbine speeds,
`respectively; T2, engine inlet temperature; and P3, compressor exit
`
`pressure. The control output from a throttling valve is the metered
`engine fuel flow rate, WFE.
`
`Th
`
`Figure 1. T55-L-712 Turboshatt Engine
`
`The control also provides a pneumatic signal to the compressor
`air bleed actuator. This signal opens and closes the compressor air
`bleed for surge protection during rapid engine transients, including
`starts. Bleed actuation is a function of an internal control parameter
`and requires no additional pilot inputs or feedback parameters.
`The T55-L-712 was qualified for production in the mid-I980's
`and has been in service since then. Information on the engine for use
`in building the simulations was obtained from a variety of sources.
`Because of the engine's age, the only digital computer simulation of
`the engine available before the current effort was a manufacturer's
`steady-state performance model. This computer model is calibrated to
`give the minimum guaranteed performance for a production engine.
`Other useful information included a design report for the engine, hy-
`dromechanical control development reports, and data from a compres-
`sor rig test. Information extracted from these sources and adapted for
`simulation use will be discussed for each model below.
`
`MODEL DESCRIPTIONS
`Two simulations of the T55 were developed, a transient simula-
`tion and a dynamic simulation. The transient T55 simulation is an
`application of version 3.0 of the Advanced Turbine Engine Simulation
`Technique (ATEST). The dynamic T55 simulation is the prototype
`application for a new dynamic turbine engine model developed at
`AEDC, version 1.0 of the Aerodynamic Turbine Engine Code
`(ATEC). These models and their application to the T55 are discussed
`below.
`The simulations were both developed in FORTRAN 77 on an
`IBM-compatible personal computer (PC). The Microsoft ° Powersta-
`tion' Version 1.0a FORTRAN development platform was used, and
`the simulations use Mi croso ft ° FORTRAN namelist files for input.
`Namelist input is not a FORTRAN 77 standard, but most current
`FORTRAN compilers have a namelist extension, so porting the simu-
`lations to other platforms is not difficult. Both simulations have been
`ported to one or more UNIX workstations, including Silicon Graphics,
`DEC and Hewlett Packard.
`Execution of either simulation on a PC or workstation requires
`only seconds or minutes of processor time. Exact execution time re-
`quired for either simulation is dependent on the computer platform
`
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`being used and the simulation length and type, as well as simulation
`user-selected options such as simulation output frequency. A repre-
`sentative run-time ratio (computer execution time / engine simulation
`time) for an ATEST T55 simulation providing output every 25 milli-
`seconds on a 90 mHz Pentium" PC with 32 Mbytes of memory is
`approximately 12:1. This means a 5 second engine acceleration would
`require about a minute to execute. A similar ATEC simulation would
`have a run-time ratio of approximately 15:1. Simulation of a dynamic
`event with ATEC would increase this due to smaller time steps and
`higher output frequency.
`
`ATEST Model
`ATEST is a modular one-dimensional component-level transient
`turbine engine model developed at AEDC (Chappell and Blevins,
`1986). ATEST 3.0 is capable of continuously simulating the full
`range of turbine engine operation, including starts, windmill, and op-
`eration from idle to maximum power. The behavior of the engine
`components is ATEST is represented by empirical data in the form of
`performance maps. The component map representations used in
`ATEST 3.0 are selected to avoid discontinuities or indeterminate val-
`ues in the start or windmill regimes. For example, the compressor
`maps use temperature ratio instead of efficiency. since at zero shaft
`speed temperature ratio is identically one, while efficiency is unde-
`fined. A detailed description of ATEST 3.0, its modeling technique,
`component map representations, and capabilities were reported by
`Chappell and McLaughlin, 1993.
`The objective of the T55 ATEST model is to simulate steady-
`state and transient operation of the engine, including fuel control,
`from startup through the normal operating region to maximum power.
`The 155 ATEST model uses six modular ATEST components: inlet,.
`compressor, burner, two turbines (gas generator and power), and exit
`diffuser (Fig. 2). Engine component maps covering operation from
`idle to maximum power were extracted from the engine manufacturer's
`steady state performance digital computer model. These maps in-
`cluded burner, combustor, gas generator turbine and power turbine
`maps, and an overall compressor map, and have been incorporated
`into the ATEST simulation. The compressor map has surge bleed
`effects imbedded in it for low speed steady-state operation.
`
`IN-ET COMPRESOR
`
`Tu:e2 OFFUSEFI
`
`SHAR2
`
`Figure 2. T55-L-712 ATEST Components
`
`The compressor maps extracted from the steady-state model were
`in the classical form of two dependent parameters, compressor pres-
`sure ratio and efficiency, as a function of two dependent parameters,
`inlet corrected airflow and corrected rotor speed. The corrected air-
`flow map was used directly; the efficiency map was converted to a
`map of compressor temperature ratio as a function of the same inde-
`pendent parameters (Fig. 3). This conversion was made using the
`
`assumption of constant ratio of specific heats, y, and specific heat at
`constamt pressure, Cp (values for air were used); the simulation's
`compressor model compensates output values from the map for actual
`gas properties.
`
`XN
`
`MTh-
`
`P3
`P2
`
`T3
`T2
`
`WA-Z/5
`
`Figure 3. Typical Compressor Performance Specification
`Curves
`
`As with the compressor, the turbine maps from the steady-state
`model required conversion for use in ATEST. The original maps had
`the form of turbine mass flow function and efficiency as a function of
`pressure ratio and corrected shaft speed. The ATEST form after con-
`version was turbine mass flow function as a function of turbine work
`factor and corrected speed, and temperature ratio as a function of pres-
`sure ratio and corrected speed (Fig. 4). The conversion was made
`using the assumption of constant y and Cp (values for air were used);
`the simulation's turbine model compensates output values from the
`map for actual gas properties.
`The ATEST burner model uses maps for pressure losses, burner
`efficiency, and combustor lightoff boundary. The pressure loss maps
`are in the form of a dry duct pressure loss and a heat addition pressure
`loss. The dry duct pressure loss is a function of corrected flow or any
`user-selected parameter, and heat addition pressure losses can be ei-
`ther calculated or tabulated. The combustion efficiency map may be
`as a function of either temperature rise and pressure, or fuel-air ratio
`and pressure. The lightoff boundary map is in the form of fuel-air
`ratio as a function of a user-defined loading parameter. The 155
`ATEST simulation uses a calculation for combined (dry and heat ad-
`dition) pressure losses, and the dry loss map is set to zero. The 155
`ATEST combustion efficiency map is tabulated as a function of fuel-
`air ratio and pressure. The pressure loss calculation and the efficiency
`map are from the manufacturer's steady-state model. The 155 ATEST
`lightoff boundary map is in the form of fuel-air ratio as a function of a
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`= HP / (XN*I) (cid:9)
`
`(I)
`
`where a is the shaft rotational acceleration, HP is the available horse-
`power, XN is the shaft speed and I is the rotor polar moment of
`inertia. Instead, the change in rotational speed is given by Eq. (2):
`
`1
`
`do) (cid:9)
`dt (cid:9)
`
`(rt — cc — rp — ro + rs) (cid:9)
`
`(2)
`
`where, w is the shaft rotational speed, Tt is the torque produced by the
`turbine, Tc is the torque required by the compressor, r p is the torque
`required to satisfy any customer power requirements, ro is the torque
`required to account for user-defined viscous or parasitic losses, and T s
`is the net torque produced by the starter and delivered to the rotor.
`155 shaft inertias were obtained from an engine design report.
`A T55 control simulation was coded from the logic block dia-
`gram shown in Fig. 6 and digitized values from the graphic schedules
`provided. This information was extracted from a nonlinear analysis
`report of the complete helicopter control system at limited flight con-
`ditions. The nonlinear analysis was to determine the transient re-
`sponse of the system at high power, so low power information on
`control schedules was limited. Although the hydromechanical control
`is analog, the digital simulation assumes a fixed control cycle time
`equal to the engine simulation transient time step of 10 milliseconds.
`The two user inputs required are Alpha] and Alpha2 as described in
`the previous section, while the output is WFE, engine fuel flow rate.
`The control simulation includes code for control actuation of the en-
`gine surge bleed, but information on the surge bleed port flow char-
`acteristics has not been included into the engine simulation. This is
`because the compressor maps already have bleed effects embedded in
`them. Use of the control signal will require the modification of the
`simulation to use separate compressor maps before and after the bleed
`port.
`
`Cr nes
`
`Combined Air Loading Factor developed by Herbert, 1957. Since no
`information was available for the T55, Herbert's flammability data for
`a generic can type combustor is used, as shown if Fig. 5.
`
`wonc Done Factor h4 - hs
`
`1.0
`
`1.0
`
`Lii
`Pt5
`
`Figure 4. Typical Turbine Performance Specification
`Curves
`
`'
`0
`5
`9. 0
`,0 0
`5
`o
`7
`5
`O. o-
`9
`0
`2
`4
`I.
`6
`I.
`
`ln
`
`5
`5
`0
`so
`
`2.
`2
`3.
`
` .
`
`. (cid:9)
`
`.
`
`..
`
`• .
`
`.
`
`.
`
`..•
`
`. •
`
`•
`
`•• (cid:9)
`
`
`
` •• (cid:9)
`
`•
`
`e (cid:9)
`
`• •
`
`: Irj
`
`Inn (cid:9)
`V
`Pd
`
`'an
`V f t./sec
`) (cid:9)
`P (cid:9)
`d ft.
`
`(
`
`son
`
`moo
`
`20
`
`0
`
`r. 305°K
`
`Figure 5. Flammablity Data as Given By Herbert, 1957
`
`Rotor dynamics during engine transient operation are modeled in
`ATEST using shaft inertias and a shaft work balance. The model uses
`a summation of torques on the shaft from attached components (e.g.
`compressors, turbines, starter, power extraction shafts) to calculate the
`change in shaft rotational speed. Torque is used instead of power
`because acceleration based on power is indeterminate at zero shaft
`speed, as shown in Eq. (1).
`
`Figure 6. Hamilton Standard JFC31-22 Hydromechanical
`Control Logic
`
`Basic steady-state and transient capability is available from the
`155 ATEST model. The model can be run with the control, requiring
`inputs of Alphal and Alpha2, or without the control, requiring inputs
`of fuel flow rate, along with the desired power turbine shaft speed.
`The model provides formatted tabular output of selected component
`and engine aerodynamic and performance data.
`
`4
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`AXIAL DISTANCE
`Figure 8. 755 Control Volumes for ATEC
`
`Because of the lack of stage data for the turbine, ATEC was re-
`stricted to modeling the turbine overall performance, not stage-by-
`stage. This allowed the use of the turbine maps developed for the
`ATEST simulation to obtain source terms for ATEC. Note that al-
`though the simulations' turbine models differ, they can use the same
`maps without modification. Multiple control volumes were specified
`in the ATEC turbine geometry to more accurately model the system
`dynamics (Garrard, 1995). The temperature and pressure changes
`across the turbine are calculated from the maps and distributed line-
`arly over the control volumes of each turbine.
`
`Reverse Flow I
`
`Zero Flow
`
`-0.2
`
`0.0 (cid:9)
`
`0.2
`
`0.4
`
`0.6
`
`Flow Coefficient
`
`0.4
`
`02
`
`-0.2
`
`Rotating_ 1 (cid:9)
`Normal
`Star rio- Operation
`1 (cid:9)
`(Pre-Stall)
`
`Discontinuity
`
`Reverse Flow
`
`Zero Flow
`
`Pressure Coefficient
`
`Temperature Coefficient
`
`-0.2 (cid:9)
`
`0.0 (cid:9)
`0.2 (cid:9)
`0.4 (cid:9)
`Flow Coefficient
`Figure 9. Typical Compressor Pressure and Temperature
`Stage Characteristics
`
`0.6
`
`ATEC Model
`ATEC is a dynamic turbine engine simulation that solves the one-
`dimensional, time dependent Euler equations with turbomachinery
`source terms. The simulation uses elemental control volumes at the
`sub-component level, such as a section of inlet duct or a compressor
`stage.
`Initial development of the ATEC model was performed at AEDC,
`as reported by Garrard. et al., 1995. The simulation is based upon the
`dynamic compressor model and simulation DYNTECC (Hale and
`Davis, 1992), which models dynamic compressor behavior at constant
`rotor speed. using an explicit solver. ATEC 1.0 has the added capa-
`bility to model rotor dynamics and employs an implicit and an explicit
`equation solver. This means engine transient simulations can be con-
`ducted efficiently with the implicit solver and its relatively large time
`steps, while maintaining the capability to simulate dynamic events
`such as compressor stall or engine surge, using the greater efficiency
`of the explicit solver at small time step sizes (Garrard, 1995).
`ATEC incorporates a unique variable time step algorithm to take
`advantage of both solvers. The algorithm uses derivative limits on
`selected variables to interactively choose the appropriate solver and
`set the solution time step. Details of the algorithm and its implemen-
`tation are reported in Garrard, 1995.
`The geometry for the model was obtained from a T55 Compres-
`sor rig and a 155 Design manual. The reverse flow combustor section
`of the engine cannot be modeled exactly in a one dimensional model,
`but the linear distance and volume of the section has been matched to
`ensure appropriate dynamic response, as presented in Fig. 7. The lack
`of as-built geometry and the modifications for the I -D model intro-
`duce recognized errors that will require trimming of the simulation
`during validation. The model consists of 45 control volumes - 12 in
`the compressor. 6 in the combustor, 3 in the gas generator turbine, 4 in
`the power turbine, and the rest in ducting (Fig. 8).
`ATEC uses a stage-by-stage model of the compressor system, as
`opposed to the lumped-component overall representation of ATEST.
`Steady-state compressor stage characteristics provide the necessary
`source terms to solve the governing equations. These characteristics
`are in the general form of a pressure coefficient and a temperature
`coefficient for each stage as a function of a flow coefficient and cor-
`rected shaft speed. as shown in Fig. 9. The actual form of these cha-
`raracteristics and their use in the dynamic model are reported in detail
`by Garrard, 1995. For the T55 simulation, the compressor stage char-
`acteristics were obtained from experimental data from a compressor
`rig test (Owen and Bobula, 1994).
`
`1
`
`0.8
`
`0.8
`Radius
`0.4
`
`0.2
`
`0
`0 (cid:9)
`-1 (cid:9)
`-2 (cid:9)
`4 (cid:9)
`3 (cid:9)
`2 (cid:9)
`1 (cid:9)
`5
`Figure 7. 755 Geometry Representation for ATEC (Garrard,
`1995)
`
`5
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`In the ATEC combustor model, the heat release source term in
`the governing energy equation is determined by the fuel energy con-
`tent, whether the combustor is lit, and, if it is burning, how well it is
`burning. The ATEC lightoff boundary is determined from a polyno-
`rnial curve fit of the data used in the ATEST map. Combustion effi- (cid:9)
`ciency is determined by using steady state engineering correlations (cid:9)
`developed by Lefebvre, 1985. Lefebvre assumed that the overall
`combustion efficiency is limited by the efficiency of fuel evaporation (cid:9)
`and the reaction efficiency. (cid:9)
`The ATEC model, like the ATEST model, accepts fuel flow rate
`as a function of time. The fuel flow rate can be input as an array or
`calculated using the control simulation described above. Since the
`control simulation uses a constant cycle time step and ATEC has a
`variable time step, an interface for the control was developed to call
`the control simulation at the appropriate time during a solution. The
`control simulation is executed at the start of the simulation run, and
`the control outputs are assumed constant until the sum of the subse-
`quent ATEC time steps is equal to or greater than the control time
`step. At this point, the control simulation is executed again with the
`current feedback variable values, and the control output values are
`updated. This interface eliminates problems with time constants in the (cid:9)
`control simulation and increases the efficiency of the simulation by (cid:9)
`avoiding unnecessary calculations during the solution.
`
`MODEL OPERATION
`The ATEST and ATEC simulations were exercised after devel-
`opment to ensure proper operation, and calibrated to available T55
`data. Details of the verification and calibration processes are dis-
`cussed below.
`
`ATEST Model
`The ATEST model has been run using parametric combinations
`of inputs in the steady-state mode, and the results were compared to
`output from the manufacturer's engine steady state performance digital
`computer model. Results for steady-state engine operation at selected
`flight conditions are shown in Figs. 10 and 11. The manufacturer's
`data are from execution of the manufacturer's model in the operating
`line mode; the ATEST data were obtained using input values for fuel
`flow rate and power turbine speed from the manufacturer's data. Fig-
`ure 10 shows a comparison between the ATEST and manufacturer's
`models of gas generator operation, and Fig. 11 shows a comparison of
`power turbine operation. The steady-state ATEST model data agree
`with the manufacturer's model data within 4 percent at sea level static
`conditions, and within 7 percent at 4000 ft, Mach 0 for high power
`operation. At low power operation, the agreement is somewhat worse.
`Transient operation of the ATEST Model has been demonstrated
`over a limited range of engine operation. The control simulation is
`limited to part power and above due to the limited control schedules
`available. Extrapolation of control schedules would extend the con-
`trol simulation operating range. Operation without the control is
`hampered by the lack of transient data to provide realistic power tur-
`bine speed inputs, so transient operation has been limited to throttle
`transients at constant power turbine speed.
`The transient ATEST results at steady input conditions are the
`same as the corresponding steady-state ATEST results. Transient
`engine data were not yet available to validate the transient ATEST
`results for transient inputs, so these results were examined qualita-
`tively for correct engine response, and were compared to manufac-
`turer's steady-state values at the transient endpoints. Representative
`
`transient inputs were assumed based on normal helicopter operation
`and pilot control inputs.
`
`1.2
`
`-
`
`.
`
`8 0.8 -
`
`
`
`ix as -
`
`.e Oa/ -
`
`8 0.2 -
`
`O
`
`0.2 (cid:9)
`
`04 (cid:9)
`
`0.6 (cid:9)
`
`0.8 (cid:9)
`
`1 (cid:9)
`
`1.2 (cid:9)
`
`IA
`
`NomxclIzod Encino uel Flow
`
`a. Sea Level Static, Standard Day
`
`1.2
`
`0.
`
`
`
`CC 0 .6
`
`f, 04
`
`8 z 0.2
`
`0.2 (cid:9)
`
`0.6 (cid:9)
`0.4 (cid:9)
`ormollz•d t noln• F u•I F low
`
`0.6
`
`b. 4000 feet, Mach 0, Hot Day
`Figure 10. ATEST and Lycoming Steady State Models Gas
`Generator Comparison
`
`0 (cid:9)
`
`0.2 (cid:9)
`
`1 (cid:9)
`0.8 (cid:9)
`0.6 (cid:9)
`0.4 (cid:9)
`Nomtlized E Vow Fuel Flow
`
`1.2 (cid:9)
`
`1.4
`
`a. Sea Level Static, Standard Day
`
`1.2
`1
`•
`▪ -
`01 3 0.8
`a
`g. as
`04
`I
`0.2
`0
`
`E
`
`0
`
`0.2 (cid:9)
`
`0.4 (cid:9)
`04 (cid:9)
`Nornxiized E rvjne Fuel Flow
`
`0.8
`
`b. 4000 feet, Mach 0, Hot Day
`Figure 11. ATEST and Lycoming Steady State Models
`Power Turbine Comparison
`
`6
`
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`
`GE-1022.006
`
`

`
`•
`
`and decreases of Alpha2 between 73 deg and 45 deg in 0.8 sec. Re- (cid:9) ▪
`
`Throttle transients between maximum continuous power and
`flight idle power at sea level static standard day conditions were (cid:9)
`
`simulated with the ATEST model. The inputs were constant Alphal (cid:9)
`of 100 deg, constant power turbine speed of 16000 rpm, and increases (cid:9)
`
`e.
`0.
`
`•
`
`ATEST
`
`TEC
`
`I.o
`
`0.6
`
`suits of the simulation are shown in Figure 12, along with the ATEST (cid:9)
`and manufacturer's model steady-state power hooks for comparison. (cid:9)
`
`• MIST
`• 4•4•1•••
`
`1-
`: 0.4
`
`7' 0.2
`cc
`
`0.0
`
`E•gl4e
`1417
`
`-1 (cid:9)
`1 (cid:9)
`2 (cid:9)
`Relative Axial Distance
`Figure 13. Steady-State Calibration of ATEC to ATEST
`Results for a Sixty Five Degree Throttle Angle Test Case
`(Garrard, 1995)
`
`3 (cid:9)
`
`4
`
`Dynamic operation of the ATEC model was operationally veri-
`fied with a test case simulating rapid engine throttle movements be-
`tween 100 percent rotor speed and 90 percent rotor speed at sea level
`static. The relative fuel flow rate input for the test case is shown in
`Fig. 14. The engine acceleration caused by the fuel flow rate increase
`was rapid enough to force the compressor into surge cycles, as shown
`by the relative compressor pressure ratio in Fig. 15. A complete dis-
`cussion of the ATEC dynamic verification is given by Garrard, 1995.
`No dynamic T55 engine data are yet available for calibration of the
`ATEC simulation. .
`1.10
`
`1.00
`
`0.90
`
`0.80
`
`Relative Fuel Flow Rate
`
`0.70
`00 (cid:9)
`
`0.5 (cid:9)
`
`1.0 (cid:9)
`
`2.5 (cid:9)
`2.0 (cid:9)
`1.5 (cid:9)
`Time (Second.)
`Figure 14. Variation of Fuel Flow Rate During Full Engine
`Operational Verification Test Case (Garrard, 1995)
`1.0
`
`3.0 (cid:9)
`
`3.5 (cid:9)
`
`4.0
`
`0.0
`00 0.5 1.0 1.5 2.0 2.5 3.0 as 4.0
`Time (Seconds)
`Figure 15. Compressor Pressure Ratio as a Function of
`Time for Full Engine Operational Verification Test Case
`(Garrard, 1995)
`
`7
`
`t (cid:9)
`
`I (cid:9)
`
`,
`
`Corrected Perow
`a. Compressor
`
`o ATETIT
`
`• 1.3011•10
`
`Engine Fuel Flow
`
`b. Power Turbine
`Figure 12. ATEST Transient and Lycoming Models
`Power Turbine Comparison
`
`ATEC Model
`The steady operation of the ATEC model and simulation was
`calibrated to the ATEST model using sea level static test casts with
`the engine operating at various throttle positions. Steady-state and
`transient ATEC results were compared to ATEST results for rases that
`were run to the same fuel flow rate. Selected results from these test
`cases are presented here, and further comparisons are detailed in Gar-
`nut 1995.
`Figure 13 gives a comparison of ATEST and ATEC simulation
`total pressures through the engine at a part power throttle position.
`The results of the effort at 100 percent thronle are compared to
`ATEST output for the same operating conditions in Table I. Com-
`paring the results to ATEST calculations, differences of less than 7
`percent were obtained and are probably due to differences in the com-
`pressor and burner component performance maps.
`
`Downloaded From: http://173.254.190.160/ on 04/13/2016 Terms of Use: http://www.asme.org/about-asme/terms-of-use
`
`GE-1022.007
`
`

`
`Location
`
`Table 1. Comparison of ATEC Performance Results to the ATEST Model
`Total Pressure (P/Pref)
`Total TemperatureiT/Tref)
`ATEST
`ATEC % Delta
`ATEST
`ATEC % Delta
`0.29
`0.29
`0.00
`0.26
`0.26
`0.00
`2.38
`2.36
`0.73
`0.51
`0.51
`0.72
`2.25
`1.76
`2.29
`1.16
`1.15
`0.47
`
`Inlet
`Compressor Exit
`Burner Exit
`Gas Generator
`Turbine Exit
`Power Generator
`Turbine Exit
`
`0.81
`
`0.87
`
`-6.27
`
`0.30
`
`0.30
`
`-0.75
`
`0.93
`
`0.75
`
`0.96
`
`-3.63
`
`0.79
`
`-5.22
`
`Lef byre, A. H. "Fuel Effects on Gas Turbine Combustion - Ig-
`nition, Stability, and Combustion Efficiency," Journal of Engineering
`for Gas Turbines and Power. Vol. 107, January 19