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עשר שנייםה העיון יום
גז וטורבינות סילון במנועי
12th Israeli Symposium on Jet
Engines
and Gas Turbines November 7 2013, Tehnion, Haifa, Israel
BOOK OF ABSTRACTS
(,3011/ – 0011) 11/7////3, ד"תשע כסלו' ד', ה יום
המעבדה למנועי סילון וטורבינות גז הפקולטה להנדסת אוירונוטיקה וחלל
חיפהן, הטכניו
http://jet-engine-lab.technion.ac.il
ענף הנעה המחלקה לאוירונוטיקה
היחידה לתשתיות-היחידה למו"פ מנהלת פיתוח אמל"ח ותשתיות
משרד הביטחון
ענף הנעה מחלקת מטוסים
להק ציוד חיל האויר
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(, 161' מס בניין) נאמן סדמו, פורשהיימר בניין, באטלר אודיטוריום
חיפה, טכניון
עשר שנייםה העיון יום
גז וטורבינות סילון במנועי
12th Israeli Symposium on Jet Engines
and Gas Turbines
November 7 2013, Tehnion, Haifa, Israel
BOOK OF ABSTRACTS
(,3011/ – 0011) 11/7////3, ד"תשע כסלו' ד', ה יום
(, 161' מס בניין) נאמן מוסד, פורשהיימר בניין, באטלר אודיטוריום
חיפה, טכניון
המעבדה למנועי סילון וטורבינות גז הפקולטה להנדסת אוירונוטיקה וחלל
ן, חיפהיוהטכנ
http://jet-engine-lab.technion.ac.il
ענף הנעה המחלקה לאוירונוטיקה
היחידה לתשתיות-היחידה למו"פ מנהלת פיתוח אמל"ח ותשתיות
משרד הביטחון
ענף הנעה מחלקת מטוסים
להק ציוד חיל האויר
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דברים לזיכרו של יהודה )אודי( גילאי ז"ל 3102כנס הנעה סילונית בטכניון; נובמבר
אודי גילאי היה מחלוצי פיתוח מערכות ההנעה מתקדמות בארץ ועסק בתחום מערכות ההנעה
. מאז צבר ניסיון מקצועי בלתי 0991 -בהנדסה אווירונאוטית בטכניון במאז סיים בהצטיינות את לימודיו
רגיל במגוון תפקידים בתחום ההנעה בכלל ומנועי סילון בפרט, בחיל האוויר, בתעשייה האוויריות וגולת
בניהול פרויקטלי והנדסי של משימות מו"פ במנועי סילון ומערכות נילוות עבור מפא"ת -הכותרת
. ומערכת הבטחון
בכל תפקידיו, גילה אודי יוזמה, מקוריות ותושייה וידע לעמוד על עקרונותיו ולהיאבק על רעיונות
מוצלחים תוך הקפדה על יושרה וחברות. בהתנהלותו המקצועית והאישית, ידע לשלב יחדיו רמות
שונות של פעילות:
טכנולוגיים חדשים, מרמת תת מערכת ועד לשותפות בעיצוב החל מיזום רעיונות וכיוונים
החזון וכיוונים לתכנית רב שנתית כוללת;
עבור דרך ליווי והנחיית בדיקות היתכנות ופרויקטי מחקר ופיתוח של מערכות הנעה שלמות
ומערכות העזר שלהם;
דע האנושיים וכלה בעיצוב ופיתוח התשתית המקצועית בארץ, בתשתיות פיזיות ובמוקדי הי
כאחד.
שנים, אך רוחו הטובה, תרומתו 8אודי היה חבר קרוב ושותף לעבודה בחברת עדמטק כמעט
ועצתו עמדו לרשותנו גם בהמשך הדרך, לאחר שקיבל על עצמו תפקיד מורכב ואתגרי ברשות התעופה
האזרחית.
השותפים פטירתו בטרם עת הינה מכה קשה למשפחתו, לחבריו ועמיתיו הרבים, ולכלל
בקהיליית העוסקים במערכות הנעה במשרד הבטחון, באקדמיה ובתעשייה. אנו שותפים לאבל
ותחושת האובדן של המשפחה. הלך מאיתנו אדם שתרומתו, אופיו ויכולותיו יחסרו לנו מאוד במעלה
הדרך.
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תודותACKNOWLEDGEMENTS
:העיון יום בקיום שתמכו ולמוסדות לגופים להודות ברצוננו
האוויר חיל
ת"מפא
טכנולוגי מכון – טכניון
לישראל
ל"רפא
הכנס לפרסום ינותודת
והחלל התעופה למדעי לאגודה
0 בישראל
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- 6 - H=Hebrew, E= English
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Lecture #A1 Behind the Curtain:
Design and Manufacturing Technology
By Alan Epstein VP Technology and Environment, Pratt & Whitney
East Hartford, CT 06108, USA The new geared engine architecture of the Pratt & Whitney PurePower® engines going into service starting next year has upended the world of narrow body commercial aviation. The design technologies which have been brought together to realize this architecture have enabled ultra-high bypass ratio propulsion to move from a vision to a reality. An equally important part of the advancement in propulsion is the development of new and improved manufacturing technology which serves as a foundation of the modern aircraft gas turbine. This lecture explores the interplay of design and manufacturing, and their implications for the future of commercial propulsion.
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Lecture #A2
Aviation Propulsion Technologies in the 21st Century
Mike Epstein Military Propulsion Engineering.
GE Aviation One of the the biggest challenges facing the aviation industry is the cost of fuel, which impacts operations and recapitalization in the commercial and military sectors. To address this issue, GE is focused on a range of near, medium, and long term strategies which features technology development, demonstration, and validation as a central element. This includes piece part, materials, and systems integration. For example, because high temperature CMC’s have only 1/3 the density of nickel alloys, a components developed using this material system will not only weigh 67% less, but potentially reduce fuel consumption by greatly reducing parasitic cooling flows. Demonstrator programs are an essential element of technology development and have directly led to much of today’s GE product line, delivering the fundamentals of materials, aerodynamics and engine architecture. Today we are working on the next generation demonstrators, including AATE, FATE, and ADVENT, which have or will run core engine and full engine demos in the next several years. These programs are validating next-generation technologies such as new materials, improved system technologies such as thermal management, new architecture ideas such as variable geometry feature in the fan and core systems, as well as improving basic aerodynamics. Finally, traditional fuel burn reduction strategies are beginning to yield diminishing returns. Innovative concepts are required that include light weight, high propulsive efficiency, highly integrated features. Also several non-Brayton cycles will be reviewed along with their benefits and challenges. Finally, aviation alternative fuels will play an increasing role in our energy future and the landscape of current and future activity will be discussed.
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Lecture # A3
Advanced Film Cooling and Combustor Cooling Concepts for Gas Turbines
Srinath V. Ekkad
Commonwealth Professor
Mechanical Engineering Department
Virginia Tech
Blacksburg VA 24060
USA
540-231-7192
In an attempt to enhance cooling efficiency, new cooling hole designs have been investigated.
Three different cooling designs are proposed: Trenched holes where the cylindrical holes are
embedded in 2-dimensional trenches to simulate slot exits; Cratered holes where the
cylindrical holes are embedded in 3-dimensional craters to reduce upward momentum; and
lastly the anti-vortex geometry where the main holes also feed two side holes to generate
anti-vortices that reduces jet lift-off and improve cooling effectiveness. All the above designs
have been tested on a flat plate in a low speed wind tunnel. Geometrical variations such as
trench width and depth, crater depth and crater-to-hole exit location, anti-vortex pair hole
size and location have been investigated.
Additional studies on Endwall contouring, combustor liner heat transfer and liner cooling
geometries will also be presented. Future directions for the gas turbine research at Virginia
Tech include the UTRC Large Scale Rotor Rig (LSRR), a rotational internal flow test rig, low
speed cascade.
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Lecture # A4
UNSTEADY AERO-THERMAL EFFECTS IN TRANSONIC TURBINES
Guillermo Paniagua
von Karman Institute, Chaussee de Waterloo 72, 1640 Rhode Saint Genese, Belgium, [email protected], http://www.vki.ac.be
In the quest for enhanced reliability in compact gas turbine engines considerable research efforts have been focused on the study of highly loaded turbines. The aero-thermo-structural performance of highly loaded designs is abated by the unsteady impact of the vane shocks on the rotor. This lecture presents the physical analysis of the stator–rotor-stator interactions in transonic turbine stages based on experimental and numerical research. The experimental assessment of the turbine was performed in unique short-duration wind tunnels, which required ad-hoc instrumentation and data processing methodologies. The detailed comparison of computational fluid dynamic results and experiments led to the understanding of the complex unsteady physics in highly loaded turbines. The figure below on the left displays the density gradient at 25% of the span at four consecutive instants, which exhibits complex vane-rotor shock interactions. The vane shock impingement on the rotor originates a separation bubble that is then tracked on the right in the entropy contour. This vortical structure is responsible for the generation of large pressure losses.
High-pressure and low-pressure turbines are mounted relatively close to each other. Therefore one needs to consider their mutual interaction. In this regard two types of research were conducted: clocking effects and the analysis of multi-splitter designs. Downstream of a turbine stage, the steady flow field is not uniform in the pitch-wise direction, clocking is the relative pitch-wise position between consecutive vanes or co-rotating rotors. The literature on clocking is related to four aspects, efficiency improvements, attenuation of the forcing in the vanes and rotor, heat transfer effects, and the transport of turbulent structures. Our research highlighted that a clocking position allowed reducing the unsteadiness in the second stator while enhancing the performance. The figure below shows the steady pressure and Nusselt number distribution in the multi-splittered low-pressure vane, investigated downstream of the high-pressure turbine. The steady heat flux data was compared to existing
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correlations for ducted flows. Local values along pressure side and suction side of passages A, B and D are presented in function of the axial coordinate. In the front section of the passage a steep acceleration occurs for both the suction and pressure sides. Along passage A, a sudden increase of the Nusselt number in the pressure side at an axial location equal to 0.15, indicating transition onset with further turbulent boundary layer development, caused by the flow deceleration associated with the diffusion along the passage. The unsteady data demonstrated the ability of the lagers airfoils to attenuate the rotor fluctuations within smaller vane passages. The present research should provide directions for the design of multi-splittered airfoils for turbine applications.
Based on the presented research, passive and active control techniques were investigated. A differential evolution algorithm was applied to optimize the transonic vane using a cost-effective Reynolds-averaged Navier–Stokes solver, computing the downstream pressure distortion and aerodynamic efficiency. Attenuation above 60% of the unsteady forcing on the rotor (downstream of the optimal vane) was observed, with no stage-efficiency abatement. On the other hand a novel proposal to control the resulting fish tail shock waves was developed based on pulsating coolant blowing through the trailing edge of the airfoils. A linear cascade representative of modern turbine bladings was specifically designed and constructed. Minimum shock intensities were achieved using pulsating cooling. The impact of cooling on wake unsteadiness for various Mach and Reynolds numbers are quantified in terms of the Strouhal number. The potential implementation of the proposed cooling scheme in turbine applications might lead to improvements on turbine efficiency and life-span.
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Lecture # A5
AHEAD –Advanced Hybrid Engines for Aircraft Development
Prof. Arvind Rao,
Technical University - Delft, The Netherlands,
Introduction
Commercial aviation has made substantial progress since its inception and is now the backbone of a modern society. In the past ten years, passenger numbers have grown by 45% and freight traffic has increased by more than 80% on a tonne-kilometre basis [1]. Moreover, aircraft emissions have reduced significantly over the last 40 years, for example, noise has reduced by 20 decibels, fuel consumption by 70%, carbon monoxide emissions by 50% and unburned hydrocarbon and smoke by 90% [1].
Although all these positive developments, aviation is facing some serious challenges towards the environment, the community and the availability of fuel. The total emission from the aviation sector is still increasing rapidly and now the sector is an active contributor to global warming. In view of these challenges, the Advisory Council for Aviation Research and Innovation in Europe (ACARE) has put forth ambitious goals for European civil aviation. The major goals being reduction of CO2 emission by 75%, NOx emission by 90% and the perceived noise levels by half in the year 2050, as compared to the baseline year 2000 [2]. To achieve these ambitious objectives, a combined improvement in the aircraft, power-plant and the air traffic management system is required.
As a consequence, the AHEAD project looks at alternative cryogenic fuels such as Liquid Hydrogen
(LH2) and Liquid Natural Gas (LNG), which seem to be most promising alternative fuels to achieve
ACARE goal of CO2 emission reduction. Due to the much higher energy density than kerosene, using
hydrogen can reduce the amount of fuel which needs to be carried onboard, however, storing such
cryogenic fuels is difficult due to the large volume requirement of such fuel. Also since these fuels have
to be stored in insulated cylindrical tanks, these fuels cannot be stored in the wings of an aircraft like
the conventional Jet-A fuel.
The Multi-Fuel Blended Wing Body Aircraft
A novel way to overcome the cryogenic fuel storage problems in aircraft is a multi-fuel Blended Wing
Body (BWB) aircraft. The wings of a BWB have sufficient room for storing LH2 tanks, without
interfering with the passenger section. Further away from the central line where wing thickness is
reduced, liquid biofuel can be stored. Such an unique configuration not only addresses the fuel
storage issues but also optimizes the usage of space in a BWB aircraft and improves the ride quality for
passengers (as passengers are located closer to the center axis of aircraft).
Schematic of the Multi Fuel Blended Wing Body Aircraft
Cryogenic Fuel
tanks
Bio fuel storage
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The Hybrid Engine
To power the novel multi-fuel BWB aircraft, a new type of propulsion system—called the hybrid
engine—has been conceived, which is able to meet the requirements of the multi-fuel BWB aircraft.
The novel engine proposed is quite different than a conventional turbofan and includes many
breakthrough technologies as listed below
Boundary Layer Ingestion (BLI): this is a method of increasing the propulsive efficiency of the
engine by embedding the engine within the airframe such that the engine can ingest the low
velocity boundary layer flow of the aircraft, reducing the engine ram drag.
Counter Rotating Fans (CRF): The proposed hybrid engine with counter rotating fans has a
smaller diameter and higher propulsive efficiency for the same bypass ratio. Furthermore,
since each stage of the fan is less loaded than a single stage fan, a CRF can sustain more non-
uniformities in the flow generated due to BLI compared to a conventional single stage fan.
Bleed Cooling The cryogenic fuel used in the proposed hybrid engine is an excellent heat sink
which can be used for cooling the bleed air, therefore, reducing the amount of bleed air
required for cooling the turbine vanes and blades. This reduces the fuel consumption of the
engine further.
The Hybrid Dual Combustion System: The proposed innovative hybrid engine uses two
combustion chambers as shown in Figure 3. The main combustor operates on hydrogen while
the second combustor (between HPT and LPT) uses biofuel in the flameless combustion mode.
Such a novel combustion system has never been used before for aero-engines. The first
combustion chamber is designed for CO2 and soot reduction, whereas the second combustion
chamber is designed for NOx reduction. Thus, reducing the harmful emissions from this engine
drastically.
Schematic of the Hybrid Engine.
Environmental impact
The use of cryogenic fuels (like Hydrogen) will have an effect on the water content in the engine exhaust. Already it was established that the proposed engine will produce contrails at a lower altitude than conventional aircraft. In the tropics where the tropopause is higher and temperatures at flight levels are commonly too high for contrail formation, the proposed Multifuel BWB aircraft will produce contrails more frequently. In order to reduce the contrail formation from the aircraft, a detailed study is being carried out by DLR
to understand the effect of soot on the contrail formation process, its optical thickness and lifetime at
various altitudes and latitudes. Coupled with this understanding, the AHEAD consortium is trying
different mission profile for the BWB aircraft for reducing the overall effect of this novel aircraft on
global warming.
- 14 -
Results and Conclusions
The “AHEAD” BWB aircraft is an environmentally friendly aircraft burning cryogenic fuels (like
LNG\LH2) and biofuels. The Multi-fuel BWB has been designed for carrying around 300 passengers for
a range of 14000 km. The comparison of the layout between the BWB and Boeing 777-200ER is
provided in Figure 4. The shorter and wider body of the aircraft makes it aerodynamically more
efficient than a conventional cylindrical body aircraft. Combined with the advanced hybrid engine, the
LNG version of the multi-fuel BWB is able to reduce CO2 emission by around 65% than a conventional
Boeing 777-200ER aircraft. For the LH2 version, the aircraft CO2 emission reduction is even more,
however a life cycle analysis of LH2 production has not been conducted yet to provide the overall CO2
reduction.
Layout comparison between the AHEAD BWB and Boeing B 777.
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Lecture # A6
"Turbojet engine compressor surge margin estimation: methods & comparison to turbojet engine experimental testing"
Prof. PRz, Dr hab. inż. Stanisław Antas,
The Faculty of Mechanical Engineering and Aeronautics, Rzeszow University of Technology, [email protected] mgr inż.
Andrzej Ćwik WSK „PZL Rzeszów S.A.”, [email protected]
The operational ratings of an aviation turbojet engine are bounded by the surge limit and choke lines on the compressor performance map. The surge line separates area of stable and unstable compressor operation. Unstable operation results in repeating fluctuations of air flow consisting on accumulation and downstream flow to the engine inlet, as a result, pressure and mass flow rate change violently. For flight conditions, crossing the surge line can result in a catastrophic loss of the thrust, dangerous aerodynamic excitations to the turbomachinery and overheating of the engine components.
At manufacturing conditions, the value of compressor surge margin KC) of the turbojet engine is determined during acceptance tests according to the criteria of a compression ratio on the steady-state operating line of a compressor (C) and a compressor turbine (CT) to the value of a compressor inlet corrected flow rate – Gl cor. Frequent cases of a necessity to change the value of surge margin of compressor thus to place the working line C-CT on the full compressor
map (forced by failing to comply with the requirement of a proper value of corG1
*
C criterion) and limited publications on
these problems caused a necessity to find an own method enabling to find basic factors controlling the value of KC at manufacturing conditions. Presentation describes conditions analysis of compressor-turbine matching for aircraft turbojet engine and theoretical research of factors affecting this matching. Recommendations are presented for rational design actions to enable change of the steady-state operating line position for compressor and gasifier turbine on the full compressor map. Influence of the throat area change of gasifier turbine nozzle guide vanes (FGV)CT and the throat area change of nozzle (FNT) on the parameters and engine compressor surge margin were measured at manufacturing conditions with the value
of c*/Glcor ratio. Results obtained from experimental tests for nozzle throat area changes verified this theoretical research. Influence of the throat area change of compressor turbine nozzle guide vanes and the throat area change of exhaust nozzle on the
engine compressor surge margin measured at manufacturing conditions had been defined with the value of corC G1
* /
ratio.
For the case when inappropriate value of the criterion (i.e., surge margin KC) is obtained during engine acceptance test, engine rebuild is requires as well as verification of the throat areas of compressor turbine nozzle guide vanes (FGV)CT and engine exhaust nozzle F5. The analytical procedure of finding the compressor surge margin function for the changes of the throat areas of the compressor turbine nozzle guide vanes and exhaust nozzle is described and presented.
Fig. 1. Control planes designations of turbojet engine with parameters changes.
*Csl Csl *
C
G1cor
(G1cor)sl
- 16 -
Fig. 2. Illustration of determining a compressor surge margin: 1 – surge limit, 2 – C-CT operating line, 3 – corrected speed curve nC cor = idem
Quantitative estimation of the distance between steady-state operating line and the surge line for chosen corrected speed curve nC cor = idem can be done by means of a compressor stable running factor determined by the formula [1]:
cor
slcor
CG
G
1
1K*
C
*
slC
(1)
where Gl cor and (Gl cor)sl is an air corrected flow rate (mass stream) determined at the compressor inlet at the working point and at the surge line suitably for nC cor = idem.
and the surge margin of compressor KC is defined by:
1 CC KK (2)
or
[%],1001% CC KK (3)
At manufacturing conditions, the value of compressor surge margin KC of the turbojet engine can be determined by the change of compressor turbine nozzle guide vanes (FGV)CT and the throat area change of exhaust
nozzle F5 . Presented definition of th analytical method of calculating the surge margin of the compressor, may be used for super-critical, critical, and sub-critical pressure ratios in engine turbine stages, i.e., when Laval number of absolute velocity at
outlet nozzle diaphragm .9,0CLλ
The comparison of values )/( 1
*
corC G obtained by analytical method and values experimentally found is presented
for the case nCcor = idem.
1
1
1'
1
*'
1
*
CTGV
CTGVcor
C
cor
C
F
FGG
(3)
'
5
5
'
1
*
1
*
1
1'
F
FGG cor
C
cor
C (4
The usage of relations (3) and (4) is very convenient but the increase of accuracy of determining changes of
)/( 1
*
corC G ratio may be easily obtained by adding correction factors.
CTGVCT
realcor
C FkeqG
)()3.(1
*
(5)
55
1
*
)4.( FkeqG
realcor
C
(6)
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Lecture # B1
Operation Principles of a Closed Loop Continuous and Heated Micro High Pressure Turbine Facility
Cukurel, Beni, Technion-Israel Institute of Technology, Lecturer at Aerospace Engineering, Haifa,
Israel. Several developmental projects are to be structured around a versatile closed-loop high
speed turbine facility, which will provide unique research capabilities to the European and Israeli research environment. A continuously running micro turbine facility will be constructed, mainly financed by the Start-up Grant provided by Technion-IIT. It will incorporate an interchangeable test section to provide hot (~600K) transonic conditions for rotating high pressure turbine stages of micro-engines. Capable of matching and varying all similarity conditions for a 150kW micro gas turbine engine, the final specifications are: turbine diameter up to 200 mm, closed loop turbine pressure ratio up to 6:1, maximum mass flow rate of 1 kg/sec, transonic blade Mach number distribution, and rotational rate of 120,000 rpm.
The power coefficient of a model turbine stage is considered “similar” to the engine representative conditions, as long as corrected mass flow, corrected speed and Reynolds
number are matched, . In an alternate notation, the total-to-total
pressure ratio can be fixed instead of the mass flow parameter, . Although the above analysis captures the pure aerodynamic performance of the turbine stages, to conduct heat transfer investigations, additional variables should be taken into consideration. By conducting energy balance at the solid fluid interface, as long as the solid to fluid thermal conductivity ratio (K) is preserved by material selection, then, Nu = f( Re, , To3/Ts). This implies that for preserving and modeling heat transfer characteristics of the gas
turbine, in addition to the purely aerodynamic variables, it suffices to retain the gas to solid temperature ratio To3/Ts.
The unique facility is developed in a manner which allows component testing of small scale
high pressure turbines, Figures 1. The main components comprising the closed test facility include a compressor, small pressure equalization tank to damp out transients, pressure drop valves, electric heater, test turbine, large dump tank and cooler, which reduces the inlet temperature to the compressor. The facility operates through the drive of a 500hp (370 kW) screw type compressor which creates pressure ratios up to 9:1, with a maximum flow rate of 1 kg/sec. In general, although the centrifugal compressors have better performance at full load, the screw compressor’s operational principles allow more versatility at reduced rotational speeds and pressure ratios. Due to better off-design operation of the screw compressor, the attainable pressure ratio and mass flow rate combinations are much larger than those of a comparable size centrifugal compressor. Moreover, in screw type compressors, since the mass flow rate is largely independent of pressure ratio, the mass flow rate of the system can be
Figure 1: Schematic of Continuous, Closed and Heated High Pressure Turbine Facility
- 18 -
well-defined by the rotational speed of the compressor shaft, independent of the pressure drop mechanisms present in the system. Finally, bypassing the after cooler located at the exit of the screw compressor provides heated flow conditions, where temperatures can reach up to 500K.
Passing through a 5m3 tank volume, which facilitates the stability of the system in transients,
the flow is diverted to a set of parallel gate valves. The valve assembly is designed such that one of the parallel arms of the system control the amount of mass flow rate diverted into the high pressure drop route. The loss mechanism is adjusted, by the control of valve closure, to the desired turbine upstream total pressure value. Therefore, the operating pressure ratio point of the compressor is decoupled from the test turbine. Further downstream, the flow passes through a 150 kW electrical mesh heater, where the downstream total temperature can be maintained up to 600 K. At the turbine, work is extracted, and the flow is expands, which reduces the total pressure and temperature. At the exit of the turbine, the flow is discharged into a large 20m3 dump tank, which maintains stable turbine exit pressure conditions. The purpose of the large volume of air is to minimize oscillatory fluctuations which can be caused due to the compressor-turbine coupling; as maintaining a slow changing quasi-steady pressure intermediate stage allows well-characterized conditions for the turbine exit and compressor inlet. Finally, before entering back into the compressor, the flow is guided through a cooler which reduces the compressor inlet flow temperature to below 320K, which is critical to prevent mechanical failures through reduced clearances. Once the desired compressor temperature is achieved, the cooler operates by principals of conservation of energy, to prevent further temperature rise in the system.
Considering the aerodynamic similarity parameters of the test turbine stage, the pressure ratio is set by combination of compressor operating point and the valve pressure loss mechanisms and the mass flow rate is predominantly defined by the compressor rotational speed. The initial mass introduced into the isolated system defines the Reynolds number, which is coarsely adjusted prior to the startup, and fine-tuned during operation. The closed loop nature of the facility allows independent aerodynamic testing of the micro gas turbine in engine representative conditions, independent from the other operational parameters including the engine speed line.
The only remaining aerodynamic parameter is the control of the turbine rotational speed. There are various ways of loading the turbine and controlling the shaft speed, such as a hydraulic-break, an AC regenerative dynamometer, or an eddy current dynamometer. The water break systems are cost prohibitive, difficult to control the load, and require an auxiliary water supply in either open loop or closed loop with cooling configurations. The AC regenerative dynamometers provide the flexibility to be used as a motor and either dissipates energy via a resistance bank or, if regulated, they can re-introduce the energy back into the grid. However, an air cooled eddy current dynamometer is significantly cheaper than either one of the alternatives, in addition to the effortless control of the rotational speed and lack of necessity for any auxiliary infrastructure. By connecting the turbine shaft, nominally rotating at 120,000 rpm, to a reduction gear with ratio 18:1, all turbine operating points, up to maximum flow rate of 1kg/sec, can be attained. The dynamometer input parameters are calculated to be PW100% = 270kW and PW50% =190 kW at design (100%) and reduced (50%) rotational speeds, whereas the corresponding torques are = 24 N-m and = 34 N-m respectively.
Having demonstrated the possibility to independently establish all aerodynamic turbine parameters, to able to achieve engine similar heat transfer distributions, the only other variable which requires unconstrained control is the gas to solid temperature ratio To3/Ts. In the presented configuration, this is achieved via the electrical heater upstream of the tested turbine stage. Furthermore, as depicted in Figure 1, to conduct cooling studies, auxiliary coolant air, which by bypasses the heater, is introduced into the test module.
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Lecture # B2
Dynamic Sealing in gas Turbines
Dr. Rammy Shellef, President, Ettem Engineering, Israel
Conventional Sealing methods
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LECTURE #B3
Using Chemical Kinetics as a Tool for Improved Efficiency and Low Emission Jet Engine
Combustor's Design
Prof. Levy Yeshayahou,
Head Turbo and Jet Engine Laboratory, TECHNION,
The design and development of low emission aero engine and gas turbines is an extremely challenging task. It requires significant data and solution of complicated physics. The requirement to satisfy projected restrictions on emissions of NOx and CO under geometrical restriction and without compromising combustion stability and efficiency over a large range of operating conditions is challenging task. The present study describes the process adopted for the design of a novel low emission jet engine combustor and its application to a specific case of Jet engine that is based on the sequential combustion technology, (the AHEAD Framework Program FP7/2007-2013 under grant agreement n° 284636). The designed jet engine includes 2 combustors. The combustor under the present development is the second combustor, situated between the high pressure and the low pressure turbines. The first combustor is operated with hydrogen fuel. Consequently the second combustor is fed with high temperature, humid and low oxygen content oxidant. Due to the specific condition of the engine, the second combustor has to operate under very lean conditions. In order to achieve stable combustion while maintaining low NOx emission, the flameless mode of combustion was adopted. Significant chemical kinetic study was invested to simulate the chemical reaction conditions within the combustor under the specific operating conditions. The study commenced by selecting an appropriate surrogate fuel that can model the multi components jet fuel under the specific operational conditions. To study the flame properties under the specific temperature, pressure and composition of the oxidant, detailed chemical kinetics analysis within a Perfectly Stirred Reactor (PSR) was performed using the CHEMKIN code. These properties included ignition delay and evolution of temperature and emissions with residence time. The data were used to determine the combustor's main dimensions, global contours and internal flow distribution. The work was followed by CFD simulation using a reduced (but verified) chemical kinetics mechanism. Several design iterations were needed in order to improve and achieve a design that comply with the geometrical constrains and the low emission requirement. At present, a reduced scale experimental model is being designed to verify its unique internal flow pattern and combustion characteristics as well as its expected low NOx and CO emissions.
Schematics description of the internal flow distribution within the "flameless combustor.
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Lecture # C1
BET SHEMESH ENGINES LTD מ"בע שמש בית מנועי
90-0090995: פקס, 90-0090099: טלפון 00999 מיקוד שמש בית המערבי התעשייה אזור
0999 סילון יומנוע גז טורבינות עיון יום - טכניון
גז טורבינת של במדחס אבק/חול בליעת נזקי להערכת מודל - תקציר
מ"בע שמש בית מנועי, לוי אלברט
אוויר מזוהמי באזורים הפועל מדחס של ראשונות דרגות בלהבי השחיקה נזקי על הסבר יינתן בהרצאה
מדברית בסביבה ואבק חול רווית באטמוספרה או .
של התלות תתואר כן-כמו. במנוע הנבלע באוויר החול ריכוז של הערכה שיטות על הסבר יינתן בהמשך
הלהב של החומר קשיות ושל הפגיעה זווית של, החלקיקים של הפגיעה במהירות השחיקה קצב .
של רוטור להב שחיקת על ב"בארה שנערכו ניסויים על המבוססת אמפירית נוסחה תוצג דבר של בסופו
האלה בפרמטרים בתלות השחיקה נזקי של כמותית ערכהה שתאפשר הליקופטר .
שורש בין להבדלים התייחסות תוך הנפגעת הלהב בפרופיל הנזק חישוב שיטת תתואר ההרצאה בסיום
הזרימה לשפת ההתקפה שפת ובין הלהב לקצה .
Abstract of Lecture
Albert levy, Bet Shemesh Engines Ltd
A model for evaluation of erosion damage due to dust/sand ingestion of the first stage of a
gas turbine compressor.
The lecture will start with an explanation about the damage caused on the first stages
blades/vanes of a gas turbine engine, operating in an air polluted area or in discrete regions
with significant concentration of dust or sand in the air.
Methods of sand concentration measurement in the air will be discussed. Quantitative
evaluation of the functional relationship between the erosion rate and the angle of impact
and its velocity will be given.
Finally an empirical formula based on tests performed to evaluate helicopter rotor blade
erosion in flight will be given including a/m parameters influence. To conclude the lecture a
method of calculation the erosion of a compressor rotor blade as a function of section (root,
mean, tip) or profile (L.E., P.S., S.S., T.E.) will be described.
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Lecture # C2
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Lecture # C3
מכונות טורבו של ואווירודינאמי מכאני בתכן אופטימיזציה
חוסיד סבלי, מילר אוהד
רפאל, מנור, סילונית הנעה מחלקת
קונפיגורציה של אופטימאליים פרמטרים בחירת לעיתים כולל סילון מנועי של ואווירודינאמיים מכאניים רכיבים תכן
ליניארית-לא סימולציה ידי-על מתבצעת המערכת דרישותל הרכיב ביצועי של התאמה בחינת. הרכיב של מסוימת
להגיע בכדי יעילים כלים ידי-על בסימולציה אוטומציה דורש זה. משתנים פרמטרים של רבות איטרציות וביצוע
..).עלות, חוזק, משקל) יעד מטרות ומספר האפשרויות-רב במרחב אופטימאלי לפתרון
חישוב ביצוע עוד מאפשרת אינה המשולבת והשפעתם רב פרמטרים רמספ, גז טורבינת כגון מורכבות במערכות
המבוססות תוכנות שכוללת ממוחשבת האופטימיזציה בתחום התקדמות חלה אחרונות בשנים. ידני אופטימיזציה
מהנדסים די-על שימוש המאפשרת לבשלות הגיעו האלו הכלים. גנטי אלגוריתם כגון מתקדמים אלגוריתמים על
.האופטימיזציה בתחום התמחות ללא
שינוי המשלבת מערכת ויושמה פותחה) ואווירודינמיקה חוזק) גז טורבינת ברכיבי אופטימיזציה לבצע במטרה
את שמנהל modeFrontier הנקרא מתקדם אופטימיזציה כלי, ם"תיב כלי עם פרמטריזאציה ידי-על גיאומטרי
. סופיים אלמנטים אנאליזות כלי בשיתוף אופטימאלי פתרון מציאת תהליך
בחשבון לוקח אשר) גנטי אלגוריתם) בעולם ביותר המתקדמים בין מתמטי בפותרן שימוש על מתבסס הרעיון
שמתכנסת כזו בצורה אותם מנתח, מטרה פונקציות הגדרת ותחת) למשל גיאומטריה אילוצי) פרמטרים מספר
.דינאמיתאווירו -ו/או מכאנית מבחינה ביותר הטוב שהינו פתרון למציאת
אופטימאלי תכן מראה אחת דוגמה. זעירים סילון מנועי עבור השיטה של מוצלח שימוש של דוגמאות יוצגו במצגת
ושטח) U-turn) מעלות 989 -ב כיוון משנה התעלה). 9 איור ראה) חום למחליף מהטורבינה חם גז תעלת של
הפסדי מינימום עם התעלה של לקונפיגורציה להגיע הינה התכן מטרת). diffuser) ליציאה מכניסה גדל התעלה
כפי הבעיה של בגיאומטריה שולטים פרמטרים לושה. גיאומטריה אילוצי על שמירה תוך, ליציאה כניסה בין לחץ
מפעיל, בפרמטרים מכוון שינוי מבצעה) 0 איור ראה) modeFrontier -ב לאופטימיזציה מודל. 9 באיור שהוצג
.הגיאומטריה את משנה) הלחץ מפל) תוצאות בדיקת ותוך) CFD (FLUENT ותוכנת) GAMBIT) מרשת
modeFrontier -ב לאופטימיזציה מודל. 0 איור התעלה צורת של פרמטריזציה. 9 איור
פרמטרים 7 של שינוי כוללת הבעיה. סילון למנוע קפיצית טבעת של מכאני בתכן אופטימיזציה מראה נוספת דוגמה
שינוי שמבצעת המערכת). 4 איור) לתכן אילוץ מהווה ליניארית-לא תזוזה כאשר) 9 יורא ראה) זמנית-בו
- 24 -
הורצו modeFrontier בעזרת. 5 באיור הוצגה מינימאלית מסה ועם נדרשת קפיציות עם טבעת לכיוון הפרמטרים
.בזמן משמעותי חיסכון תוך הטבעת של שונות קונפיגורציות 989
הטבעת של פרמטרי ודלמ. 9 איור
הטבעת של תזוזות סימולציית. 4 איור
התהליך של השלבים כל לבין הגיאומטריים הפרמטריים בין המקשר Workflow. 5 איור
.לחוזק טורבינה להב של אופטימיזציה של נוספת דוגמה תוצג במצגת
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Lecture # D1
High-Entropy Alloys for Aerospace Applications
H. Rosenson, A. Fleisher, S. Essel and A. Katz Israel Institute of Metals
High-entropy alloys (HEA) are materials newly introduced to the research society having unique ability to precipitate solid solution phases from the melt rather than intermetallics. Usually they contain at least five metallic elements with equiatomic concentrations from 5% to 35% of each one. To encourage formation of solid solution phases, highly negative mutual mixing entropies are required. HEA's thermodynamic order is less than of conventional polycrystalline materials but higher than of amorphous materials. The majority of high-
entropy alloys is Ni-based and then mainly composed of FCC and/or BCC solid solutions. They are characterized by extremely high after-ageing hardness (HV900), wear resistance, high temperature stability and fatigue resistance, low temperature dependence of the strength and fatigue endurance limit comparable to the presently used high-temperature materials. Due to these properties they may be successfully used for turbine aircraft engine blades and vanes production instead of the presently used superalloys which require intermetallic-based
coating. The following research will be conducted around the development of industrial casting technique for these materials and products. In association with Beith-Shemesh Engines LTD and Institute National de Cercetare Dezvoltare pentru Metale Neferoase si Rare (IMNR, Romania) the research will develop the required technology for the casting of blades from HEA. The technical challenges and scientific motivation for this development are presented
alongside the description of the work planned. HEA of two various compositions have been produced by the following production route (according to the reported in the literature): arc-melting of commercial/high-purity metals mixture in high-purity argon atmosphere, remelting three times to improve homogeneity and followed by direct solidification in cold copper mold. Microstructural examination has been performed by means of SEM/EPMA and XRD. The mechanical properties (tensile testing and microhardness measurement) have been examined for comparison with presently used Ni-based superalloys. The obtained results confirmed advantages of HEA mentioned in the
literature.
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Lecture # D2
Introduction In the course of Selected Topics in Aerospace Propulsion the students are asked to group into
small teams and work together on conceptual design of a jet-engine for selected
requirements, while putting an emphasis on the hot-sections of the engine.
The teams are required to work on different aspects regarding the engine-design, relying on
topics studied in the course meetings.
Presenter's team, a group of 5 students, took the task of designing a jet-engine for a combat
UAV (such as the experimental X-47A). The engine, called SE'ARA (The Hebrew meaning for
storm), that is shown here, is a result of countless hours of calculations, meetings with Dr.
Boris Glezer (The course lecturer) and many trial-&-error loops.
The design started with basic specs and customer's required performance: Thrust of 5000 lbf
at altitude 20000 ft with Mach 1.3. Then, continued through the thermodynamic calculations,
stress analysis, heat transfer and finished with tip-clearance mechanism and also a 1:1 layout
of the entire engine.
The presenter's task in this project was the design of engine's layout- One can basically
describe it as the incorporation of all ideas and approaches, presented by team members, into
one conceptual design. The requirement of designing an engine for a combat UAV and
meeting the target thrust became one the most significant restrictions in terms of dimensions.
This was done by down-scaling known engines, as part of market research, and by trying to
follow the exact dimensions which came out of the thermodynamic calculations. Attempts to
utilize the free space within disc-cavities were carried out while applying common rules and
coefficients, such as the "1:4 chord rule" for vane-rotor-spacing. Considerations about
placement of bearings, cooling, lubricant system and assembly process were also taken in
mind. In few cases, due to our decision to use quite unorthodox approaches, some innovative
and creative thinking was required. An example of that is team's approach for dealing with
tip-clearance issues; A rather simple solution for a complex problem, suggested by Dr. Glezer,
was implemented.
The results of this work will be shown in the following presentation.
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Lecture # D3
ומנוגדות קשות מפרט דרישות עם סילון מנוע של קונספטואלי תכן 'יאקירביץ אלי: המרצה
התכנון נקודות) שונות עבודה בנקודות גבוה ודחף נמוכה דלק צריכת -" ומנוגדות קשות דרישות" הגדרת
).מאך ומספרי בגובה שונות הנדרשות תכן מרחב כולל צירים מערכת הגדרת SFC דחף מול
SFC פרמטר לכל מפרט דרישות מול מתבצע הנרמול) מנורמל דחף מול רמלמנו(
ואף ציר לכל שונות תכנון בנקודות אלא, הצירים לשני זהה עבודה בנקודת או התכנון בנקודת דווקא לאו .עקומה לכל
.בדידה -" מפוזרת" מיפוי לדיאגראמת דוגמה, קונפיגורציות למיפוי הצורך הגדרת - מסקנות כולל), ב"ארה של האוויר חיל) המיפוי נבנה שבהשראתו המאמר, צירי דו מנוע של מיפוי הצגת
.פרמטרים של השפעות" מצפן" .הקונפיגורציות מיפת על ההשפעות את שמציג תרשים בעזרת והשפעתן תכנון נקודות על הסבר
Conceptual Design of a Jet Engine with Difficult and Contradicting Design Requirements Lecturer: Eli Yakirevich
Definition of "Difficult and Contradicting Design Requirements" - low SFC and high thrust at different operating points (the operating points vary with altitude and Mach number). Definition of the coordinate systems to define the design space
SFC vs. thrust
Normalized SFC vs. normalized thrust (normalization is according to design
requirement per parameter)
Not necessarily at the design point or even at the same operating point for the different axes. Presentation of the necessity for mapping the configurations, presentation of a discrete cases diagram. Presentation of a mapping of a twin spool engine, the reference article for this sort of mapping (U.S. Air Force), including conclusions - a "compass" for the influence of the parameters on the design space. Design point(s) - explanation its influence on the design mapping and graphical presentation.
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Lecture # E1
F-4 (Phantom) A/C modernization by engines' conversion
Emanuel Liban
Chairman of the Israel Society of Mechanical and Aeronautical Engineers
PW-1120 מניפה טורבו במנועי G.E. J-79 סילון מנועי החלפת ידי על) F-4) קורנס מטוס שדרוג
ליבן עמנואל
המהנדסים לשכת, ותעופה מכונות מהנדסי אגודת ר"יו
In parallel to the development of the Lavi A/C, the Israeli Air Force and Israel Aircraft Industries
launched an upgrade program of the F-4 A/C dubbed "Phantom 2000" that consisted of new weapon
systems and the retrofit of existing J-79 jet engine with the new PW-1120 turbo fan engines which
were developed for the Lavi A/C.
In the lecture the design and modifications of the different systems will be described as well as flight
demonstrations.
The new PW-1120 engines improved the A/C performance in range and agility, increased the power to
rate ratio and enhance low altitude high speed capability.
The a/c was flown in the Paris air show in 1991 and was a major attraction.
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Lecture # E2
"Minimizing Flutter in the T700 Turbo shaft Engine"
:רקע
וינשוף) AH-64) שרף/פתן מסוקי על מורכבים GE חברת מתוצרת T700-701C/D מנועי
(UH-60 (אוויר בחיל.
בהזדקרות המתבטאת בהתנעה פרפור תופעת החלה בחיל המסוקים קליטת מתקופת החל
.המותרת המכסימלית למגבלה עד טמפרטורה בעליית המלווה ההתנעה במהלך מנוע
הגיחה זה מסוג מתקלות כתוצאה. חם המנוע כאשר שנייה בהתנעה לרוב התרחשה התופעה
.בוטלה המתוכננת
הוסרו השנה אותה במהלך, T701D לתצורת המנועים הסבת עם החמירה התופעה 0999 משנת
.הפרפור תופעת בעקבות מנועים 49 מעל
.מבצעיות גיחות לביטולי השאר בין וגרם אוויר צוותי את הטריד הנושא
:מטרה
מנוע הסרות) האחזקה עלויות את להפחית מ"ע החוזרות בהתנעות הפרפור לתופעת מענה מתן
.גיחה ביטולי צמצום י"ע המנוע אמינות ושיפור) תמתוכננו בלתי
:שבוצעה פעילות
:מישורים במספר התופעה חקירת
'.ד בדרג) מדחס) הקר החלק חקירת .9
.במנוע הדלק ווסת חקירת .0
.מנועים הרצת במתקן ניסויים ביצוע .9
.מנועים כ"בסד ויישומם מסקנות הסקת .4
:שבוצעה מהפעילות מסקנות
.הדלק ווסתיל אופטימלי כיול ביצוע .9
.ההמרה נוסחת ותיקוף יישום י"ע הספרות לדרישות הבדיקה מתקן תוצאות התאמת .0
:הפעילות תוצרי
.חוזרות בהתנעות הפרפור תופעת של משמעותי צמצום .9
.אחזקה עלויות צמצום .0
.המנוע אמינות שיפור .9
- 31 -
Lecture # E3
FAILURE OF HPT BLADES – CFM56-3 ENGINE
- 32 -
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Lecture # F1
An Update on the Application of Multivariate Statistics to Improve and Automate Engine Trending and Diagnostics (ET&D)
David Ogden, Director Susan Zubik, Manager
Southwest Research Institute® (SwRI®) SwRI has been implementing the technology presented at the 2012 Israeli Symposium on Jet Engines and Gas Turbines that included the use of multivariate statics as a core technology for detecting anomalous engine performance leading to diagnostics and repair. The challenge is to design a system that has high sensitivity to small changes in system performance, but conversely does not create a significant level of false alarms, even across a wide range of operating and environmental conditions. Previously implemented technologies for this purpose did not achieve either of these goals. If the system has low sensitivity, important changes in performance are missed and no action will result. If the system has a high false alarm rate, then the users quickly discount or ignore the results and revert to an undefined manual diagnostic process based on personal experience. Another challenge is detecting changes in the engine condition using only existing sensors. Adding new sensors to legacy systems is not cost effective and any proposals to add sensors will likely be rejected by the user. All of the examples included in this presentation were accomplished using only existing sensors. The process for analyzing the engine performance requires data collection, feature extraction, anomaly detection, and diagnostics. SwRI has been developed the feature extraction algorithms for the F100 engine and is integrating them into the existing USAF data collection infrastructure. Feature extraction requires identification of operating regimes and filtering data for noise and environmental variations. SwRI’s proprietary nSPCTTM software is integrated into the system to provide the anomaly detection function. The nSPCT outputs are being connected to the USAF maintenance and repair processes to allow the maintainers confirm and repair the degraded engine. The work that SwRI is performing will support both -220 and -229 engine variants. The process is being tested using USAF archived data and maintenance records. The 0.1% false alarm rate has been statistically confirmed. The next steps for the program include deploying the system to the field and applying it to other engine models and additional fleets in other countries.
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Lecture # F2
Maintaining Gas Turbine Compressor Efficiency
Steve Deane,
Technical Sales Manager – ZOK International Group Ltd
Gas Turbine operators are under increasing pressure to ensure that their equipment
continues to operate as efficiently and economically as possible. With as much as 60% of the
output power being consumed by the compressor assembly it is vital that gas turbine
compressor efficiency is monitored and maintained.
This presentation addresses the basic principles of compressor operation before looking in
more detail at the causes and effects of contamination on the efficiency of Airfoil surfaces,
corrosion of Airfoil surfaces and erosion of Airfoil surfaces.
We will look at the causes and types of fouling which result in performance degradation which
directly influences the efficient operation of the Gas Turbine Compressor Assembly.
We will also cover damage limitation and condition monitoring before addressing the benefits
of Compressor Washing. The presentation will close with a case study outlining how analysis
of the water wash samples can provide a valuable insight into the effectiveness of a specific
wash regime and how this information can be used to support compressor efficiency
condition monitoring data.
The presentation utilises material from extensive research conducted in Compressor Fouling,
case study examples and knowledge gained from over 30 years of experience in the
manufacture and distribution of Gas Turbine Compressor Cleaning Fluids.
- 35 -
Lecture # F3
- 36 -
Notes
- 37 -
Notes
