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Aerodynamic Investigation of Leading Edge Contouring and External Cooling on a Transonic Turbine Vane
KTH, School of Industrial Engineering and Management (ITM), Energy Technology, Heat and Power Technology. (Turbinebomachinery)ORCID iD: 0000-0001-5162-2289
2014 (English)Doctoral thesis, comprehensive summary (Other academic)
Abstract [en]

Efficiency improvement in turbomachines is an important aspect in reducing the use of fossil-based fuel and thereby reducing carbon dioxide emissions in order to achieve a sustainable future. Gas turbines are mainly fossil-based turbomachines powering aviation and land-based power plants. In line with the present situation and the vision for the future, gas turbine engines will retain their central importance in coming decades. Though the world has made significant advancements in gas turbine technology development over past few decades, there are yet many design features remaining unexplored or worth further improvement. These features might have a great potential to increase efficiency. The high pressure turbine (HPT) stage is one of the most important elements of the engine where the increased efficiency has a significant influence on the overall efficiency as downstream losses are substantially affected by the prehistory. The overall objective of the thesis is to contribute to the development of gas turbine efficiency improvements in relation to the HPT stage.

 

Hence, this study has been incorporated into a research project that investigates leading edge contouring near endwall by fillet and external cooling on a nozzle guide vane with a common goal to contribute to the development of the HPT stage. In the search for HPT stage efficiency gains, leading edge contouring near the endwall is one of the methods found in the published literature that showed a potential to increase the efficiency by decreasing the amount of secondary losses. However, more attention is necessary regarding the realistic use of the leading edge fillet. On the other hand, external cooling has a significant influence on the HPT stage efficiency and more attention is needed regarding the aerodynamic implication of the external cooling. Therefore, the aerodynamic influence of a leading edge fillet and external cooling, here film cooling at profile and endwall as well as TE cooling, on losses and flow field have been investigated in the present work. The keystone of this research project has been an experimental investigation of a modern nozzle guide vane using a transonic annular sector cascade. Detailed investigations of the annular sector cascade have been presented using a geometric replica of a three dimensional gas turbine nozzle guide vane. Results from this investigation have led to a number of new important findings and also confirmed some conclusions established in previous investigations to enhance the understanding of complex turbine flows and associated losses.

 

The experimental investigations of the leading edge contouring by fillet indicate a unique outcome which is that the leading edge fillet has no significant effect on the flow and secondary losses of the investigated nozzle guide vane. The reason why the leading edge fillet does not affect the losses is due to the use of a three-dimensional vane with an existing typical fillet over the full hub and tip profile. Findings also reveal that the complex secondary flow depends heavily on the incoming boundary layer. The investigation of the external cooling indicates that a coolant discharge leads to an increase of profile losses compared to the uncooled case. Discharges on the profile suction side and through the trailing edge slot are most prone to the increase in profile losses. Results also reveal that individual film cooling rows have a weak mutual effect. A superposition principle of these influences is followed in the midspan region. An important finding is that the discharge through the trailing edge leads to an increase in the exit flow angle in line with an increase of losses and a mixture mass flow. Results also indicate that secondary losses can be reduced by the influence of the coolant discharge. In general, the exit flow angle increases considerably in the secondary flow zone compared to the midspan zone in all cases. Regarding the cooling influence, the distinct change in exit flow angle in the area of secondary flows is not noticeable at any cooling configuration compared to the uncooled case. This interesting zone requires an additional, accurate study. The investigation of a cooled vane, using a tracer gas carbon dioxide (CO2), reveals that the upstream platform film coolant is concentrated along the suction surfaces and does not reach the pressure side of the hub surface, leaving it less protected from the hot gas. This indicates a strong interaction of the secondary flow and cooling showing that the influence of the secondary flow cannot be easily influenced.

 

The overall outcome enhances the understanding of complex turbine flows, loss behaviour of cooled blade, secondary flow and interaction of cooling and secondary flow and provides recommendations to the turbine designers regarding the leading edge contouring and external cooling. Additionally, this study has provided to a number of new significant results and a vast amount of data, especially on profile and secondary losses and exit flow angles, which are believed to be helpful for the gas turbine community and for the validation of analytical and numerical calculations.

Abstract [sv]

Ökad verkningsgrad i turbomaskiner är en viktig del i strävan att minska användningen av fossila bränslen och därmed minska växthuseffekten för att uppnå en hållbar framtid. Gasturbinen är huvudsakligen fossilbränslebaserad, och driver luftfart samt landbaserad kraftproduktion. Enligt rådande läge och framtidsutsikter bibehåller gasturbinen denna centrala roll under kommande decennier. Trots betydande framsteg inom gasturbinteknik under de senaste årtionden finns fortfarande många designaspekter kvar att utforska och vidareutveckla. Dessa designaspekter kan ha stor potential till ökad verkningsgrad. Högtrycksturbinsteget är en av de viktigaste delarna av gasturbinen, där verkningsgraden har betydande inverkan på den totala verkningsgraden eftersom förluster kraftigt påverkas av tidigare förlopp. Huvudsyftet med denna studie är att bidra till verkningsgradsförbättringar i högtrycksturbinsteget.

 

Studien är del i ett forskningsprojekt som undersöker ledskenans framkantskontur vid ändväggarna samt extern kylning, i jakten på dessa förbättringar. Den aerodynamiska inverkan av en förändrad geometri vid ledskenans ändväggar har i tidigare studier visat potential för ökad verkningsgrad genom minskade sekundärförluster. Ytterligare fokus krävs dock, med användning av en rimlig hålkälsradie. Samtidigt har extern kylning i form av filmkylning stor inverkan på verkningsgraden hos högtrycksturbinsteget och forskning behövs med fokus på den aerodynamiska inverkan. Av denna anledning studeras här inverkan både av ändrad hålkälsradie vid ledskenans framkant samt extern kylning i form av filmkylning av skovel, ändvägg och bakkant på aerodynamiska förluster och strömningsfält. Huvudpelaren i detta forskningsprojekt har varit en experimentell undersökning av en geometrisk replika av en modern tredimensionell gasturbinstator i en transonisk annulärkaskad. Detaljerade undersökningar i annulärkaskaden har gett betydande resultat, och bekräftat vissa tidigare studier. Detta har lett till ökad förståelsen för de komplexa flöden och förluster som karakteriserar gasturbiner.

 

De experimentella undersökningarna av förändrad framkantsgeometri leder till den unika slutsatsen att den modifierade hålkälsradien inte har någon betydande inverkan på strömningsfältet eller sekundärförluster av den undersökta ledskenan. Anledningen till att förändringen inte påverkar förlusterna är i detta fall den tredimensionella karaktären hos ledskenan med en redan existerande typisk framkantsgeometri. Undersökningarna visar också att de komplexa sekundärströmningarna är kraftigt beroende av det inkommande gränsskiktet. Undersökning av extern kylning visar att kylflödet leder till en ökad profilförlust. Kylflöde på sugsidan samt bakkanten har störst inverkan på profilförlusten. Resultaten visar också att individuella filmkylningsrader har liten påverkan sinsemellan och kan behandlas genom en superpositionsprincip längs mittsnittet. En viktig slutsats är att kylflöde vid bakkanten leder till ökad utloppsvinkel tillsammans med ökade förluster och massflöde. Resultat tuder på att sekundärströmning kan minskas genom ökad kylning. Generellt ökar utloppsvinkeln markant i den sekundära flödeszonen jämfört med mittsnittet för alla undersökta fall. Den kraftiga förändringen i utloppsvinkel är dock inte märkbar i den sekundära flödeszonen i något av kylfallen jämfört med de okylda referensfallet. Denna zon fordrar ytterligare studier. Spårgasundersökning av ledskenan med koldioxid (CO2) visar att plattformskylning uppströms ledskenan koncentreras till skovelns sugsida, och når inte trycksidan som därmed lämnas mer utsatt för het gas. Detta påvisar den kraftiga interaktionen mellan sekundärströmning och kylflöden, och att inverkan från sekundärströmningen ej enkelt kan påverkas.

De generella resultaten från undersökningen ökar förståelsen av komplexa turbinflöden, förlustbeteenden för kylda ledskenor, interaktionen mellan sekundärströmning och kylflöden, och ger rekommendationer för turbinkonstruktörer kring förändrad framkantsgeometri i kombination med extern kylning. Dessutom har studien gett betydande resultat och en stor mängd data, särskilt rörande profil- och sekundärförluster och utloppsvinkel, vilket tros kunna vara till stor hjälp för gasturbinssamfundet vid validering av analytiska och numeriska beräkningar.

Place, publisher, year, edition, pages
Stockholm: KTH Royal Institute of Technology, 2014. , 125 p.
Series
TRITA-KRV, ISSN 1100-7990 ; 14:04
Keyword [en]
gas turbine, aerodynamics, secondary flow, external cooling, trailing edge cooling, film cooling, aerodynamic loss, high pressure turbine, nozzle guide vane
National Category
Aerospace Engineering Applied Mechanics Energy Engineering
Research subject
Aerospace Engineering; Energy Technology
Identifiers
URN: urn:nbn:se:kth:diva-150458ISBN: 978-91-7595-240-6 (print)OAI: oai:DiVA.org:kth-150458DiVA: diva2:743500
Public defence
2014-10-01, Kollegiesalen, Brinellvägen 8, KTH, Stockholm, 10:00 (English)
Opponent
Supervisors
Projects
Turbopower, Sector rig
Note

QC 20140909

Available from: 2014-09-09 Created: 2014-09-04 Last updated: 2014-10-13Bibliographically approved
List of papers
1. Experimental studies of leading edge contouring influence on secondary losses in transonic turbines
Open this publication in new window or tab >>Experimental studies of leading edge contouring influence on secondary losses in transonic turbines
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2012 (English)In: ASME Turbo Expo 2012: Turbine Technical Conference and Exposition, ASME Press, 2012, 1109-1119 p.Conference paper, Published paper (Refereed)
Abstract [en]

An experimental study of the hub leading edge contouring using fillets is performed in an annular sector cascade to observe the influence of secondary flows and aerodynamic losses. The investigated vane is a three dimensional gas turbine guide vane (geometrically similar) with a mid-span aspect ratio of 0.46. The measurements are carried out on the leading edge fillet and baseline cases using pneumatic probes. Significant precautions have been taken to increase the accuracy of the measurements. The investigations are performed for a wide range of operating exit Mach numbers from 0.5 to 0.9 at a design inlet flow angle of 90°. Data presented include the loading, fields of total pressures, exit flow angles, radial flow angles, as well as profile and secondary losses. The vane has a small profile loss of approximately 2.5 % and secondary loss of about 1.1%. Contour plots of vorticity distributions and velocity vectors indicate there is a small influence of the vortex-structure in endwall regions when the leading edge fillet is used. Compared to the baseline case the loss for the filleted case is lower up to 13 % of span and higher from 13% to 20 % of the span for a reference condition with Mach no. of 0.9. For the filleted case, there is a small increase of turning up to 15 % of the span and then a small decrease up to 35 % of the span. Hence, there are no significant influences on the losses and turning for the filleted case. Results lead to the conclusion that one cannot expect a noticeable effect of leading edge contouring on the aerodynamic efficiency for the investigated 1st stage vane of a modern gas turbine.

Place, publisher, year, edition, pages
ASME Press, 2012
Series
Proceedings of the ASME Turbo Expo, 8
Keyword
Aerodynamic efficiency, Aerodynamic loss, Experimental studies, Reference condition, Transonic turbine, Turbine guide vane, Velocity vectors, Vorticity distribution
National Category
Energy Engineering
Identifiers
urn:nbn:se:kth:diva-92183 (URN)10.1115/GT2012-68497 (DOI)000335720900101 ()2-s2.0-84881178157 (Scopus ID)978-079184474-8 (ISBN)
Conference
ASME Turbo Expo 2012: Turbine Technical Conference and Exposition, GT 2012; Copenhagen; Denmark; 11 June 2012 through 15 June 2012
Note

QC 20130115

Available from: 2012-03-28 Created: 2012-03-28 Last updated: 2014-10-09Bibliographically approved
2. Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade
Open this publication in new window or tab >>Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade
2012 (English)In: Proceedings of the ASME Turbo Expo, ASME Press, 2012Conference paper, Published paper (Refereed)
Abstract [en]

An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream.

Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.

Place, publisher, year, edition, pages
ASME Press, 2012
National Category
Energy Engineering
Identifiers
urn:nbn:se:kth:diva-92186 (URN)10.1115/GT2012-68088 (DOI)2-s2.0-84881132398 (Scopus ID)978-079184474-8 (ISBN)
Conference
ASME Turbo Expo 2012 - Turbine Technical Conference and Exposition, Copenhagen, June 11-15, 2012
Note

QC 20150706

Available from: 2012-03-28 Created: 2012-03-28 Last updated: 2015-07-06Bibliographically approved
3. Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade
Open this publication in new window or tab >>Suction and Pressure Side Film Cooling Influence on Vane Aero Performance in a Transonic Annular Cascade
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2013 (English)In: Proceedings of the ASME Turbo Expo, 2013Conference paper, Published paper (Refereed)
Abstract [en]

An experimental study on a film cooled nozzle guide vane has been conducted in a transonic annular sector to observe the influence of suction and pressure side film cooling on aerodynamic performance. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at an exit reference Mach number of 0.89. The aerodynamic results using a five hole miniature probe are quantified and compared with the baseline case which is uncooled. Results lead to a conclusion that the aerodynamic loss is influenced substantially with the change of the cooling flow rate regardless the positions of the cooling rows. The aerodynamic loss is very sensitive to the blowing ratio and a value of blowing ratio higher than one leads to a considerable higher loss penalty. The suction side film cooling has larger influence on the aerodynamic loss compared to the pressure side film cooling. Pitch-averaged exit flow angles around midspan remain unaffected at moderate blowing ratio. The secondary loss decreases (greater decrease in the tip region compared to the hub region) with inserting cooling air for all cases compared to the uncooled case.

Series
Proceedings of the ASME Turbo Expo, Vol. 6 A
Keyword
Aero-dynamic performance, Aerodynamic loss, Blowing ratio, Engine components, High pressure gas, Nozzle guide vanes, Secondary loss, Transonic annular cascade
National Category
Engineering and Technology Mechanical Engineering Aerospace Engineering Applied Mechanics Energy Engineering
Research subject
SRA - Energy
Identifiers
urn:nbn:se:kth:diva-124520 (URN)10.1115/GT2013-94319 (DOI)2-s2.0-84890151073 (Scopus ID)978-079185522-5 (ISBN)
Conference
ASME Turbo Expo 2013 Turbine Technical Conference and Exposition, GT 2013; San Antonio, Tx, United States, 3-7 June, 2013
Note

QC 20140626

Available from: 2013-07-08 Created: 2013-07-08 Last updated: 2014-09-09Bibliographically approved
4. Aerodynamic Implication of Endwall and Profile Film Cooling in a Transonic Annular Cascade
Open this publication in new window or tab >>Aerodynamic Implication of Endwall and Profile Film Cooling in a Transonic Annular Cascade
2013 (English)In: 21st ISABE Conference / [ed] ISABE, Busan, Korea, 2013Conference paper, Published paper (Refereed)
Abstract [en]

An experimental study is performed to observe the aerodynamic implications of endwall and profile film cooling on flow structures and aerodynamic losses. The investigated vane is a geometrically similar transonic nozzle guide vane with engine-representative cooling geometry. Furthermore, a new formulation of the cooling aerodynamic loss equation is presented and compared with the conventional methods. Results from a 5-hole pneumatic probe show that the film coolant significantly alters the secondary flow structure. The effect of different assumptions for the loss calculation is shown to significantly change the measured loss.

Place, publisher, year, edition, pages
Busan, Korea: , 2013
Series
ISABE-2013, ISABE-2013-1155
National Category
Engineering and Technology Mechanical Engineering Aerospace Engineering Energy Engineering Applied Mechanics
Research subject
SRA - Energy
Identifiers
urn:nbn:se:kth:diva-128934 (URN)
Conference
21st International Symposium on Air Breathing Engines (ISABE-2103), September 9-13, Busan, Korea
Projects
Sector Rig
Note

QC 20140409

Available from: 2013-09-17 Created: 2013-09-17 Last updated: 2014-09-09Bibliographically approved
5. Influence of Prehistory and Leading Edge Contouring on Aero Performance of a Three-Dimensional Nozzle Guide Vane
Open this publication in new window or tab >>Influence of Prehistory and Leading Edge Contouring on Aero Performance of a Three-Dimensional Nozzle Guide Vane
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2014 (English)In: Journal of turbomachinery, ISSN 0889-504X, E-ISSN 1528-8900, Vol. 136, no 7, 071014-1-071014-10 p.Article in journal (Refereed) Published
Abstract [en]

Experiments are conducted to investigate the effect of the prehistory in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (five hole and three hole) concentrating mainly on the end wall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector, and vorticity contour, as well as mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the prehistory (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface, which allows identifying the locations of secondary flow vortices, stagnation line, and saddle point.

Place, publisher, year, edition, pages
ASME Press, 2014
Keyword
Secondary Flows, Blade Passage, Fillets
National Category
Aerospace Engineering Applied Mechanics Energy Engineering
Research subject
SRA - Energy
Identifiers
urn:nbn:se:kth:diva-139946 (URN)10.1115/1.4026076 (DOI)000335964100014 ()2-s2.0-84994246462 (Scopus ID)
Funder
Swedish Energy Agency
Note

QC 20140612

Available from: 2014-01-15 Created: 2014-01-15 Last updated: 2017-12-06Bibliographically approved
6. Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance
Open this publication in new window or tab >>Shower Head and Trailing Edge Cooling Influence on Transonic Vane Aero Performance
Show others...
2014 (English)In: Journal of turbomachinery, ISSN 0889-504X, E-ISSN 1528-8900, Vol. 136, no 11, 111001- p.Article in journal (Refereed) Published
Abstract [en]

An experimental investigation on a cooled nozzle guide vane (NGV) has been conducted in an annular sector to quantify aerodynamic influences of shower head (SH) and trailing edge (TE) cooling. The investigated vane is a typical high pressure gas turbine vane, geometrically similar to a real engine component, operated at a reference exit Mach number of 0.89. The investigations have been performed for various coolant-to-mainstream mass-flux ratios. New loss equations are derived and implemented regarding coolant aerodynamic losses. Results lead to a conclusion that both TE cooling and SH film cooling increase the aerodynamic loss compared to an uncooled case. In addition, the TE cooling has higher aerodynamic loss compared to the SH cooling. Secondary losses decrease with inserting SH film cooling compared to the uncooled case. The TE cooling appears to have less impact on the secondary loss compared to the SH cooling. Area-averaged exit flow angles around midspan increase for the TE cooling.

Place, publisher, year, edition, pages
ASME Press, 2014
Keyword
Aerodynamic loss, shower head cooling, trailing edge cooling, film cooling, nozzle guide vane
National Category
Aerospace Engineering Applied Mechanics Energy Engineering
Research subject
Aerospace Engineering; Energy Technology
Identifiers
urn:nbn:se:kth:diva-148421 (URN)10.1115/1.4028024 (DOI)000343933500001 ()2-s2.0-84994276021 (Scopus ID)
Projects
Sector rig
Note

QC 20140912

Available from: 2014-08-07 Created: 2014-08-07 Last updated: 2017-12-05Bibliographically approved
7. Aerodynamic Investigation of External Cooling and Applicability of Superposition
Open this publication in new window or tab >>Aerodynamic Investigation of External Cooling and Applicability of Superposition
2015 (English)In: 11th EUROPEAN CONFERENCE ON TURBOMACHINERY FLUID DYNAMICS AND THERMODYNAMICS, EUROPEAN TURBOMACHINERY SOC-EUROTURBO , 2015Conference paper, Published paper (Other academic)
Abstract [en]

An experimental investigation of the overall external cooling on a cooled nozzle guide vanehas been conducted in a transonic annular sector cascade. The investigated vane is a typicaltransonic high pressure gas turbine vane, geometrically similar to a real engine component.The investigations are performed for various coolant-to-mainstream mass-flux ratios. Resultsindicate that the aerodynamic loss is influenced substantially with the change of the coolingflow. Area-averaged exit flow angles in midspan region are unaffected at moderate filmcoolant flows, for all cooling configurations except for trailing edge cooling. The trailing edgecooling decreases the turning in all investigated cases. Results lead to a conclusion that bothtrailing edge and suction side cooling have significant influence on the aerodynamic losswhereas the shower head cooling is less sensitive to the loss. Investigations with individualcooling features essentially lead to the applicability of the superposition technique regardingthe aerodynamic loss for film cooled vanes, which is this paper’s contribution to the researchfield. Results show that the superposition technique can be used for the profile loss but not forthe secondary loss.

Place, publisher, year, edition, pages
EUROPEAN TURBOMACHINERY SOC-EUROTURBO, 2015
Keyword
Aerodynamic loss, film cooling, trailing edge cooling, nozzle guide vane, superposition
National Category
Aerospace Engineering Applied Mechanics Energy Engineering
Research subject
Aerospace Engineering; Energy Technology
Identifiers
urn:nbn:se:kth:diva-150455 (URN)000380606100056 ()2-s2.0-84983171152 (Scopus ID)
Conference
11th EUROPEAN CONFERENCE ON TURBOMACHINERY FLUID DYNAMICS AND THERMODYNAMICS, Madrid Spain, MAR 23-26, 2015
Note

QC 20161111

Available from: 2014-09-04 Created: 2014-09-04 Last updated: 2016-11-11Bibliographically approved

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