An Ultra-High Bypass Ratio Turbofan Engine for the Future

Transcription

1 Undergraduate Team Engine Student Design Competition 2014/15 An Ultra-High Bypass Ratio Turbofan Engine for the Future - Request for Proposal - September 13, 2014

2 2 Abstract Major engine manufacturers are continually assessing and revising their technical & business plans to ensure that their vision reaches into the next decade. In the commercial aviation market, replacement engines for new generations of the Boeing 787 and Airbus A380 & A350 airplanes are currently being considered. Very recently Rolls-Royce revealed its road map for the future (*), whereby it will extend its Trent 1000 and Trent XWB engine programs to address significantly higher bypass ratios, further improvements in propulsive efficiencies at cruise and reduced fuel burn & emissions for long range travel in 2025 and beyond. This Request for Proposal asks that you also look to 2025 and design a new geared, 3-spool, high bypass ratio turbofan for entry into service around that time for use on twin-engine, wide-body passenger and freight aircraft. Your primary objective is also reduced fuel burn, as a result of higher propulsive efficiency at cruise conditions. A generic model, representative of 3-spool current systems is supplied as a baseline engine. This model has been generated solely on the basis of publically-available information. You should model this engine with your design system to provide a viable reference from which to gauge your improvements. You are then required to retain the core of the baseline design and generate a new LP/IP system that fits around it, using aerodynamic similarity. A simple but typical, multi-segment, extended mission should be constructed that covers both design-point and offdesign engine operations. Such a mission will also test propulsive efficiencies at cruise and reduced fuel burn specifically and should be flown using both engines. The performance characteristics and total fuel consumption of both engines should be estimated over the mission and stated clearly in the proposal. The benefits of the new design should be clearly stated. Special attention should be paid to engine mass, dimensions & integration with the aircraft. Technical feasibility is critical and operating costs should also be considered. Dr. Ian Halliwell AIAA Air Breathing Propulsion Group and IGTI Aircraft Engines & Education Committees Principal Engineer, PSM-Alstom ianhalliwell@earthlink.net (*) Aviation Week and Space Technology, August 25, 2014.

3 3 CONTENTS Page 1. Introduction 4 2. Design Objectives & Requirements 6 3. Baseline Engine Model Overall Characteristics Inlet Fan Intermediate-Pressure Compressor Inter-Compressor Duct High-Pressure Compressor Combustor High-Pressure Turbine Inter-Turbine Duct Intermediate-Pressure Turbine Inter-Turbine Duct Low-Pressure Turbine Core Exhaust & Nozzle Bypass Duct Hints & Suggestions Competition Expectations 33 References 34 Suggested Reading 35 Available Software & Reference Material 35 Appendix 1. Letter of Intent 37 Appendix 2. Rules and Guidelines 38 I. General Rules 38 II. Copyright 39 III. Schedule & Activity Sequences 39 IV. Proposal Requirements 39 V. Basis for Judging 40

4 4 1. Introduction The A is the first Airbus A350 model and seats 314 passengers in a three-class cabin and 9-abreast layout. It has a standard design range target of 15,000 km (8,100 nmi). The 900 is designed to compete with the Boeing ER and replace the Airbus A A 900R variant, which has been proposed but not yet launched, would feature higher engine thrust and a strengthened structure. Figure 1: Rolls-Royce Trent XWB Engines on the Airbus A350 The range of the A R is estimated to be 17,600 km (9,500 nmi), which would be boosted to about 19,100 km (10,315 nmi) by these design improvements to compete with the Boeing LR and be capable of non-stop flight from London-Heathrow to Auckland. Rolls-Royce agreed with Airbus to supply a new variant of the Trent engine - correspondingly named the Trent XWB - for the A350 XWB aircraft. After low-speed wind tunnel tests, Airbus froze the static thrust at sea level for all three proposed variants in the kn (74,000 94,000 lbf) range in This Request for Proposal is aimed at future engines for this type of aircraft and for this type of mission.

5 5 Figures 1 and 2 are from the Airbus and Rolls-Royce websites respectively and show the A350 XWB aircraft and the installed engine. Figure 2: A Closer View of the Rolls-Royce Trent XWB Engine Installation General characteristics Capacity 314 passengers (3-class) Length m Wing span 69.8 m Height m Wing area 443 m² Max. take-off weight 268 t Power plant 2 high bypass ratio turbofans; 374 kn each at take-off Performance Cruise speed Mach 0.85 Range 17,600 km Service ceiling km Table 1: General Characteristics of the Airbus A350-XWR Aircraft

6 6 Some typical aircraft characteristics are given in Table 1. At take-off, the total thrust needed from each of the two engines is 374 kn. Table 2 contains a summary of basic engine characteristics, taken mostly from References 1 and 2. This data provides target values for the baseline engine model. It is emphasized that the model is intended to be only a rough generic representation of the Rolls-Royce Trent XWB purely for the purposes of this exercise! Design Features of the Baseline Engine Engine Type Axial, turbofan Number of fan/booster/compressor stages 1, 8, 6 Number of HP/LP turbine stages 1, 2, 7 Combustor type Annular Maximum net thrust at sea level kn Specific fuel consumption at cruise at Mach 0.85 & altitude g/kn.s Overall pressure ratio at max. power 50.0 Bypass ratio 9.3 Max. envelope diameter m Max. envelope length m Dry weight less tail-pipe 5,445 kg Table 2: Baseline Engine Design Targets: Basic Data, Overall Geometry & Performance The actual baseline engine model is described in some detail in Section Design Objectives & Requirements A new engine design is required for future versions of the Airbus A350 and Boeing 777 and 787, with an entry-into-service date of The new engine should include a geared fan. An explanation should be given for why a gear is needed, along with the pros and cons! The current flight envelope ranges from take-off at static sea-level conditions to cruise at 12,190 m/mach This is to be retained for the new engine, so these two flight conditions should be used as the principal design points for candidate engines. Maximum potential take-off thrust should match that of the baseline engine described later and the actual take-off thrust given in Table 1 should be assumed. The range target outlined in

7 7 the Introduction should be borne in mind and it is hoped that the endurance might be extended with the new derivative engine by reducing the fuel consumption and minimizing engine mass. In the baseline engine model, a nominal power off-take of only 50 kw has been assumed. This is too low and should be increased to 200 kw. The generic baseline engine model should be used as a starting point. The core of the engine (HP compressor, combustor & HP turbine) should be retained and a new LP/IP system should be designed around it. The bypass ratio should be increased to 15 by using a geared fan. The overall pressure ratio should be increased to 60. A turbine inlet temperature of 1784K has been assumed in the baseline engine. In the new design, based on the entry into service date, assume that advances in materials and cooling technology permit a T4 limit of 1930K. The development and potential application of carbon matrix composites is of particular interest. Based on research of available literature, justify carefully your choices of any new materials, their location within the engine and the appropriate advances in design limits that they provide. Aerodynamic similarity should be used to ensure compatibility between conditions at the IPC exit and HPC inlet. The new engine design should be optimized for minimum engine mass & fuel burn, based on trade studies to determine the best combination of fan & intermediate compressor pressure ratios, bypass ratio, overall pressure ratio and turbine entry temperature. Values of these four major design parameters should be compatible with those expected to be available in 2025 and the selected design limits should be justified in the proposal. For the gear, assume a mass of kg per hp transmitted. Design proposals must include engine mass, engine dimensions, net thrust values, specific fuel consumption, thermal and propulsive efficiencies at take-off (standard sealevel conditions) and cruise. Details of the major flow path components must be given. These include inlet, fan, IP Compressor, HP compressor, combustor, HP turbine, IP turbine, LP turbine, exhaust nozzle, bypass duct, and the connecting ducts. Since reduced specific fuel consumption does not necessarily lead to reduced fuel consumption should the new engine be heavier, the fuel burn over an assumed mission must be determined by dividing it into suitable segments in terms of time at altitude and Mach number and summing the incremental fuel burn estimates. This should be done for both the baseline engine and the new derivative to determine the improvement.

8 8 3. Baseline Engine Model As stated previously, the baseline engine is a 3-spool, high bypass ratio turbofan. A generic model has been generated from publically-available information (References 1 & 2) using GasTurb12. Some details of the baseline model are given below to assist with construction of your baseline case and to provide some indication of typical values of design parameters. Figure 3: An Unmixed, High Bypass Ratio Turbofan Engine Schematic with Calculation Stations & Nominal Cooling Flows Figure 3 contains a general schematic with relevant station numbers.

9 9 3.1 Overall Characteristics Major Design Parameters In a turbofan engine, the four primary design variables are turbine entry temperature (T 4 ), overall pressure ratio (OPR or P 3 /P 2 ), fan pressure ratio (FPR or P 21 /P 2 ) and bypass ratio (BPR). We usually differentiate between the fan pressure ratios in the core & bypass streams. In a 3-spool engine an additional variable is introduced in the form of pressure ratio generated by the intermediate pressure compressor, so an optimum IPC/HPC pressure ratio split must therefore be determined. Table 3: Basic Input Table 3 is the Basic Input for the GasTurb12 model of the baseline engine and the five primary design variables are specified. To generate an acceptable replica of the engine, a unique combination of the remainder must be estimated iteratively using performance figures which are provided namely the net thrust (F N ) and specific fuel consumption (sfc) at cruises conditions - as targets. Since the sfc target is not at the engine design point, this can only be checked

10 10 periodically once what is thought to be a satisfactory design point solution has first been obtained. Table 3 also contains some of the secondary inputs, while the remainder are addressed below. The first row of Table 3 assumes negligible total pressure loss between the inlet leading edge and the fan face. The inner and outer fan pressure ratios are then selected separately; there is more blade speed at the fan tip than at its hub, so the inner & outer fan pressure ratios have been set at 1.4 & 1.43 respectively fairly aggressive but not unreasonable for a modern single-stage machine. A zero total pressure loss is then accounted for in the duct between the fan and the IP compressor or booster. This is probably optimistic but not too much so, as the prevailing Mach number is quite low. Knowing that the required overall pressure ratio is 50.0, results in a pressure ratio across the remainder of the compression system of , allowing for losses. This is distributed between the booster and the HP compressor with 6.3 across the former (over 8 stages) and 5.76 across the latter (over 6 stages). A 2.5% total pressure loss is assumed in the bypass duct. Inter-turbine duct losses of 0.8% and zero have been used again somewhat optimistic! Continuing with the input description, the design bypass ratio was set at 9.3. A value of K for the turbine exit temperature was taken as being reasonable for this engine with limited cooling capacity and an expected long life for the HP turbine (say 5,000 hours). The temperature is a guessed value as, understandably, engine manufacturers do not reveal such critical information. In fact, this value of T 4 was the result of an iterative process that involved turbomachinery efficiencies and the target thrust. The next four parameters relate to the primary combustor; they are all fairly conventional values by modern standards. The burner part load constant is an element in the calculation of burner efficiency that is discussed in the GasTurb12 User Guide in Reference 3. Without expert knowledge, this is best left alone! The remaining parameters in Table 3 may be considered as secondary influences and are discussed briefly below. Secondary Design Parameters Cooling Air: Mention has already been made of bleed and cooling air flows the secondary flows. Only the overboard bleed is listed in Table 3 (although this is in fact zero), however the secondary flows indicated in Figure 2 have been set via another air system tab on the input screen as fractions of W 25, the HP compressor entry flow. Pressure Losses: A number of total pressure losses, mentioned earlier, are also specified in Table 3 by inserting the appropriate pressure ratios across the inter-compressor duct, the inter-turbine duct, the mixer and the primary combustor. Turbomachinery Efficiencies: Efficiencies of the fan, HP compressor, HP turbine and LP turbine are entered via their respective tabs on the input screen. The values are not listed specifically in Table 3, but may be reviewed in the output summary presented later in Table 4. The designer has the choice of either isentropic or polytropic values, so he or she should be certain of their applicability and their definitions! Both values appear in

11 11 the output summary in Table 4. However, another option is available that has been used here for both compressors & fan and turbines. It allows GasTurb12 to estimate turbine efficiencies from data supplied via values of stage loading and flow coefficients. For turbines these values are used in a Smith Chart (Reference 4), assuming an equal work spilt between stages. It is recommended that either this be used or initial values be taken from Table 4. Power Off-take: All engines have power extracted in the Trent XWB it is taken from the IP spool via a bevel gear and a tower shaft that passes through an enlarged vane or strut in the frame between the IP and HP compressors. This is often preferred to the use of a separate auxiliary power unit, depending on how much power is required for airframe use. In the application currently under consideration, considerable auxiliary power may be needed for avionics and passenger equipment and this usage is growing rapidly in modern aircraft. We have selected a nominal power off-take of 50 kw from our baseline engine but 200kW has been requested in Design Objectives in Section 2. A limited study has been made of the influence of a number of secondary parameters and it was determined that the default values present in the GasTurb12 generic model should be retained, based on the known expertise of the author of the code. Dimensions: Diameters & Lengths The engine cycle may be defined purely on the basis of thermodynamics. Geometrically, we define a rubber engine initially - where performance is delivered in terms of a net thrust of kn given in Table 4 once the engine scale has been determined. We also have a target dimensional envelope to fit into, namely a maximum casing diameter of m and length of m, although the latter is very much open to interpretation. The diameter can be determined via the mass flow rate; the length is a separate issue that is dealt with by manipulation of vane & blade aspect ratios and axial gaps in the turbomachinery and by suitable selection of duct lengths, usually defined as fractions of the corresponding entry radii. Once the correct thrust has been reached, the maximum radius is determined by setting an inlet radius ratio and then varying the Mach number at entry to the fan. These values are input on the primary input screen under the LP compressor tab, where a Mach number of was combined with a fan inlet radius ratio of and a fan tip speed of m/s were found to be appropriate. This sets the general radial dimension for the complete engine, although in fact downstream of the fan, the entry radii of the IP and HP compressors are determined independently. The HP & LP turbine radii follow from the exit values of the respective upstream components. For the ducts, radial dimensions are keyed off the inner wall with the blade spans being superimposed. For the overall engine length, early adjustments are made by eye (My personal philosophy is that if it looks right, it probably is right!), with final manipulations being added as the target dimension is approached. The fan diameter turned out to be 3.02 m (compared to the target value of m in Table 2.) while the overall diameter of the engine model is m, which allows for the thickness of the nacelle. The engine model length of m includes the exhaust system and cone so this is deemed to be satisfactory. The target length of m in Table 2 may be interpreted as a flange-to-

12 12 flange length that extends from the fan face to just downstream of the rear frame, just aft of the LP turbine in the plot or our engine model in Figure 5. The target value was in fact taken from the Trent 1000, as the Trent XWB data was not available in Reference 2. (We are not cutting metal here folks, so we are probably OK!) Materials & Weights Table 4: A Summary of the Baseline Engine Model As far as possible, use was made of the materials database in the GasTurb12 design code. For proprietary reasons many advanced materials are not included. Examples of these are: polymeric composites used in cold parts of the engine, such as the inlet and fan; metal matrix composites, which might be expected in the exhaust system; carbon-matrix-composites, again intended for use in hot sections. All of these materials are considerably lighter than conventional alternatives, although it should be noted they may not yet have found their way into the baseline engine, where long life and reliability are critical. However, within the component models, material

13 13 densities can be modified independently of the database and I have taken advantage of this feature in some cases where I believe that advanced materials of lower density are appropriate. Use has also been made of the materials data in Reference 5, interpolating and extrapolating where necessary. In GasTurb12 component weights are calculated by multiplying the effective volumes by the corresponding material densities. Of course, only the major elements which are designed directly are weighed and there are many more constituents. Nuts, bolts, washers, seals and other much larger elements such as fuel lines, oil lines, pumps and control systems still must be accounted for. In the engine industry, this is done usually, at the preliminary design stage, by the application of a multiplier or adder whose value is based on decades of experience. In general, a multiplication factor of 1.3 is recommended in the GasTurb12 manual, but for an engine as large as the Trent XWB I reduced this to a net mass factor of 1.15 in Table 5 mainly because it got me closer to the gross engine weight I was looking for! The total mass of the engine shown in Table 5 (5, kg) is 2.4% over the 5,445 kg target in Table 2, but it should be remembered that the tail pipe is not accounted for in the latter and in our model the core nozzle weighs kg when the mass factor has been applied. Conveniently, this accounts for most of the discrepancy! A summary of the baseline engine model is presented in Table 4 and Table 5 is a more detailed Overall Output Table.

14 14 Table 5: Baseline Engine Detailed Output A cutaway of the baseline engine is shown in Figure 4.

15 15 Figure 4: A Cutaway View of the Rolls-Royce Trent XWB Engine A plot of the GasTurb12 baseline engine model appears in Figure 5.

16 16 Figure 5: GasTurb12 Model of the Rolls-Royce Trent XWB - the Baseline Engine Some details of the component models now follow.

17 Inlet The inlet is designed with a conical center body (Figure 5). In practice, a single-stage fan can be cantilevered from a bearing located in the main frame of the engine. The outer diameter of the inlet has been determined from that of the fan. Table 6: Inlet Design Pertinent characteristics of the inlet are shown in Table 6. At kg, the inlet is fairly light and this is because, based on the density, we have taken a typical Ti-Al alloy as our choice of materials. It is noteworthy that the GasTurb inlet is merely the portion of the casing (plus center body) immediately upstream of the fan. The GasTurb12 model begins at the upstream flange, which is located further forward of the central cone than shown in the real engine in Figure Fan Table 7: Fan: Detailed Overview

18 18 The fan characteristics are given in Tables 7 & 8. The radius ratio and inlet Mach number are of particular interest because, when taken with mass flow rate, they define the fan tip radius. Based on tip radius, the blade tip speed sets the rotational speed of the LP spool. The value of corrected flow per unit area ( kg/m 2 or lbm/ft 2 ) is fairly conventional and corresponds to the input value of Mach number (0.574). Table 8: Fan General Output On September 12, 2014 three new parameters were added to the LPC input of GasTurb12 to control the inlet duct to the IPC. The new inputs are indicated in red in Table 8. If necessary, an update to the code should be acquired by users.

19 Intermediate-Pressure Compressor Inputs for the intermediate pressure compressor are provided in Tables 7 & 8. To maintain access to the engine geometry and plot, it may be necessary to s switch to the efficiency known option and insert the estimated isentropic value. Table 7: Intermediate Compressor - Detailed Overview Table 9: Intermediate Compressor - General Output

20 Inter-Compressor Duct Input and output for the inter-compressor duct are given in Table 10. Table 10: Inter-Compressor Duct Notice that in addition to using an overall net mass factor to adjust the engine weight, individual net mass factors may be applied to the components or net mass adders may be used, although this remains at a value of unity for the inter-compressor duct since very little of the structure is left unaccounted for in the simple model. 3.6 High Pressure Compressor Table 11: High Pressure Compressor - Detailed Overview Again, we set the speed of the HP spool via the tip speed and the corresponding radius. The general characteristics of the HP compressor are given in Table 11. Input and output parameters are shown in Table 12.

21 Table 12: High Pressure Compressor - General Output 21

22 Combustor A fairly conventional annular combustor is used and details are given in Table 13. The high density of its material corresponds to the necessary thermal properties. The combustor is a major structural component, linked closely to the HP turbine first vane assembly. Table 13: Combustor 3.8 High-Pressure Turbine Table 14: High Pressure Turbine Basis for Efficiency Estimate As stated in Section 3.1, the efficiency of the high pressure turbine was estimated by GasTurb12 on the basis of the data shown in Table 14, which is made available once that efficiency option is selected. As a result of that selection, the details of the HP turbine in Table 14 appear.

23 23 Table 15: HPT Summary A general summary of the HP turbine is given in Table 16, followed by the velocity diagrams and Smith Chart in Figure 6.

24 24 Table 16: High Pressure Turbine General Output Figure 6: High Pressure Turbine Velocity Diagrams & Smith Chart

25 Inter-Turbine Duct 1 Table 17 contains details of the inter-turbine duct between the HP and IP turbines. Its relatively short length allows the two turbines to be close-coupled and the exit-to-inlet radius ratio of 1.1 emphasizes this. The intermediate shaft rotates counter to those of the LP and HP systems although this is not indicated in the velocity diagrams shown here. Table 17: Inter-Turbine Duct Intermediate-Pressure Turbine Table 18: Intermediate Pressure Turbine Basis for Efficiency Estimate Table 18 contains the input data used when the option to calculate turbine efficiency is selected. The warning on the high exit radius ratio appears because the value is beyond conventional limits but it is due to the high value of bypass ratio and the relatively small size of the inner engine flowpath. To maintain access to the engine geometry and plot, it may be necessary later to switch to the efficiency known option and insert the calculated isentropic value.

26 26 Table 19: Intermediate Pressure Turbine Summary As a result of the efficiency calculation option, Table 19 appears in the IP turbine output. The stage loading coefficient is fairly conventional but the stage flow coefficient is quite high 1. These observations are reflected in the velocity diagram in Figure 7. The rotational speed of the IP spool was set primarily by turbine disk stress considerations, but an increase in axial velocity could have improved IP turbine performance. Additional input and output characteristics of the IP turbine are given in Table Loading coefficient (Ψ) = ΔH/U 2. Flow coefficient (Φ) = V ax /U.

27 27 Table 20: Intermediate Pressure Turbine General Output Figure 7: Intermediate Pressure Turbine Velocity Diagrams & Smith Chart

28 Inter-Turbine Duct 2 Table 21: Inter-Turbine Duct 2 Table 21 contains input and output information for the second inter-turbine duct between the IP and LP turbines. The exit/inlet radius ratio increases the radial location of the LP turbine and results in higher blade speeds, lower loading coefficients and hence improved efficiencies Low-Pressure Turbine Characteristics of the low pressure turbine are presented in Tables and Figure 8. Figure 8 contains velocity diagrams for the first and last stages. The flared nature of the LP turbine flowpath ensures that meanline radii are maximized, stage loading coefficients are minimized and stage efficiencies are fairly. However, it may be seen from Figure 8 that the common design point for all seven stages is too far to the left on the Smith Chart due mainly to the high mean blade speed and improvements in the form of higher efficiency and smaller disks could be obtained by reducing rpm. It should be noted that the efficiency contours in Figure 8 (and Figure 7 & 9) are expressed as fractions of the maximum value on the chart! The true value of the average stage efficiency is 91.89%, which corresponds to the value in the engine performance summary in Table 4. Table 22: Basis for LP Turbine Calculated Efficiency

29 Table 23: LPT Summary 29

30 30 Table 24: Low Pressure Turbine: General Output Figure 8: Low Pressure Turbine Velocity Diagrams & Smith Chart

31 Core Exhaust & Core Nozzle The core exhaust is directly downstream of the low pressure turbine. It is comprised of an outer casing, an inner casing, and an inner cone that closes off the casing, and a strut or frame. In Figure 5 on page 16, the core exhaust extends to about 5.6 m. The core exhaust in GasTurb12 does not include the convergent portion or the core nozzle. Table 25 contains the input and output details of the core exhaust while Table 26 covers the remainder, termed the core nozzle. The cone ends in the exhaust duct Table 25: Core Exhaust The core nozzle is the part of the engine that converges to its exit area at about 6.33 m in Figure 4. The casing material density in the core nozzle is the same as that for the core exhaust, although a lighter material most likely could have been used owing to the local temperatures. Table 26: Core Nozzle

32 Bypass Duct Table 27 defines the input and output parameters for the bypass duct. The shape and geometric continuity of the bypass duct with adjacent structures depends critically on the values of the parameters indicated by the blue box. Table 27: Bypass Duct 4. Hints & Suggestions You should first model the baseline engine with the same software that you will use for your new engine design. Your results may not match the generic baseline model exactly but will provide an essential starting point for a valid comparison of weights and performance for your new engine. In general, subsonic commercial engines tend to be sized at take-off rather than at topof-climb (the beginning of cruise). However, since the major objective in this exercise is to minimize fuel burn at cruise where most of the fuel will be burned it is essential that off-design performance (particularly for the turbines) be given special attention. The efficiencies of the turbomachinery components may be assumed to be the same as those of the baseline engine, and be input directly or the calculate efficiency mode of GasTurb12 may be invoked. This is not an aircraft design competition, so credit will not be given for detailed derivation of aircraft flight characteristics, but some reasonable assumption should be made - and clearly stated - concerning the thrust needed by the airplane compared to the engine capabilities at a particular Mach number and altitude.

33 33 The use of design codes from industrial or government contacts, that are not accessible to all participating teams, is not allowed. Even though the date for submission of Letters of Intent is stated as November 1, 2014 on pages 37 and 39, it is recommended that teams who know that they will enter the competition inform AIAA, ASME-IGTI or Dr. Ian Halliwell as soon as possible, so that assistance may be given and access to design codes may be arranged, where appropriate (See page 35). Questions will be taken by volunteers from the AIAA Air Breathing Propulsion Technical Group or the IGTI Aircraft Engines Technical Committee, whose contact information will be provided to teams who submit a letter of intent. 5. Competition Expectations The existing rules and guidelines for the Student Design Competition shall be observed and these are provided in Appendix 2. In addition, the following specific suggestions are offered for the event. This is a preliminary engine design. It is not expected that student teams produce design solutions of industrial quality, however it is hoped that attention will be paid to the practical difficulties encountered in a real-world design situation and that these will be recognized and acknowledged. If such difficulties can be resolved quantitatively, appropriate credit will be given. If suitable design tools and/or knowledge are not available, then a qualitative description of an approach to address the issues is quite acceptable. In a preliminary engine design the following features must be provided: Definition and justification of the mission and the critical mission point(s) that drive the candidate propulsion system design. Clear and concise demonstration that the overall engine performance satisfies the mission requirements. Documentation of the trade studies conducted to determine the preferred engine cycle parameters such as fan pressure ratio, bypass ratio, overall pressure ratio, turbine inlet temperature, etc. An engine configuration with a plot of the flow path that shows how the major components fit together, with comments on operability at different mission points.

34 34 A clear demonstration of design feasibility, with attention having been paid to technology limits. Examples of some, but not all, velocity diagrams are important to demonstrate viability of turbomachinery components. Stage count estimates, again, with attention having been paid to technology limits. Estimates of component performance and overall engine performance to show that the assumptions made in the cycle have been achieved. While only the preliminary design of major components in the engine flow path is expected to be addressed quantitatively in the proposals, it is intended that the role of secondary systems such as fuel & lubrication be given serious consideration in terms of modifications and how they would be integrated in to the new engine design. Credit will be given for clear descriptions of how any appropriate upgrades would be incorporated and how they would affect the engine cycle. Each proposal should contain a brief discussion of any computer codes or Microsoft Excel spreadsheets used to perform engine design & analysis, with emphasis on any additional special features generated by the team. Proposals should be limited to fifty pages, which will not include the administrative/contents or the signature pages. References 1. Road Map: Rolls-Royce s future turbofan strategy will leverage European, national and company research. Aviation Week & Space Technology. August 25, Aerospace Source Book. Aviation Week & Space Technology. January 26, GasTurb 12: A Design & Off-Design Performance Program for Gas Turbines < Joachim Kurzke, A Simple Correlation of Turbine Efficiency S. F. Smith Journal of the Royal Aeronautical Society. Volume Aeronautical Vest Pocket Handbook. Pratt & Whitney Aircraft. Circa 1980

35 35 Suggested Reading 1. Gas Turbine Theory H.I.H Saravanamuttoo, G.F.C Rogers &.H. Cohen, Prentice Hall. 5 th Edition Aircraft Engine Design J.D.Mattingly, W.H. Heiser, & D.H. Daley AIAA Education Series Elements of Propulsion Gas Turbines and Rockets J.D. Mattingly. AIAA Education Series Jet Propulsion N. Cumpsty. Cambridge University Press Gas Turbine Performance P. Walsh & P. Fletcher. Blackwell/ASME Press. 2 nd Edition, Fundamentals of Jet Propulsion with Applications Ronald D. Flack Cambridge University Press The Jet Engine Rolls-Royce plc Aircraft Propulsion 2 nd Edition. Saeed Farokhi. John Wiley & Sons Ltd Available Software & Additional Reference Material GasTurb 12 is a comprehensive code for the preliminary design of propulsion and industrial gas turbine engines (Reference 3). It encompasses design point and off-design performance, based on extensive libraries of engine architectures and component performance maps, all coupled to impressive graphics. A materials database and plotting capabilities enable a detailed engine model to be generated, with stressed disks and component weights. A student license for this code is available at a very low price directly from sales@gasturb.de strictly for academic work only.

36 36 AxSTREAM is the first design & analysis code that permits the topic of propulsion and power generation by gas & steam turbine to progress beyond velocity diagrams in the course of university class. A suite of compressor and turbine modules cover the design steps from meanline and streamline solutions to detailed design of airfoils. Use of this code is also supported fully by excellent graphics. SoftInWay Inc. recently announced the availability of AxSTREAM Lite to students that covers the design of turbines. However, an expanded license will be provided to participants in the Undergraduate Team Engine Design Competition that also includes fans and compressors for an appropriate time period prior to submission of proposals. Once a Letter of Intent has been received, the names of team members will be recognized as being eligible to be granted access to the AxSTREAM software. Students must then apply to SoftInWay Inc. SoftInWay will not contact team members. GSP is NLR's ( primary gas turbine performance simulation tool ( It is a component based modeling environment based on a flexible objectoriented architecture that allows modelers to simulate steady-state and transient performance of virtually any gas turbine configuration using a user-friendly drag-and-drop interface. GSP has been used for a variety of applications such as various types of off-design performance analysis, emission calculations, control system design and diagnostics of both aircraft and industrial gas turbines. All team managers or supervisors of the competing design teams are welcome to request a free team license. The offers above are subject to ITAR restrictions.

37 37 Appendix 1. Letter of Intent Undergraduate Team Engine Design Competition 2014/15 Request for Proposal: An Ultra-High Bypass Ratio Turbofan Engine for the Future Title of Design Proposal: Name of School: Designer s Name AIAA or ASME Graduation Date Degree Team Leader Team Leader AIAA Foundation will act as the administrator for this competition. In order to be eligible for the 2014/2015 Undergraduate Team Engine Design Competition, you must complete this form and return it electronically to the AIAA Student Programs Coordinator, Rachel Andino (rachela@aiaa.org) before November 1, 2014, at AIAA Headquarters, as noted in Appendix 2, Section III, Schedule and Activity Sequences. Signature of Faculty Advisor Signature of Project Advisor Date Faculty Advisor Printed Project Advisor Printed Date

38 38 Appendix 2. Rules and Guidelines I. General Rules 1. All undergraduate AIAA or ASME branch or at-large Student Members are eligible and encouraged to participate. 2. Teams will be groups of not more than four students. 3. An electronic copy of the report in MS Word or Adobe PDF format must be submitted on a CD or DVD to AIAA Student Programs. Total size of the file(s) cannot exceed 60 MB, which must also fit on 50 double spaced, 12 point font pages when printed. The file title should include the team name and/or university. A Signature page must be included in the report and indicate all participants, including faculty and project advisors, along with their AIAA or ASME member numbers. Designs that are submitted must be the work of the students, but guidance may come from the Faculty/Project Advisor and should be accurately acknowledged. Graduate student participation in any form is prohibited. 4. Design projects that are used as part of an organized classroom requirement are eligible and encouraged for competition. 5. More than one design may be submitted from students at any one school. 6. If a design group withdraws their project from the competition, the team chairman must notify AIAA Headquarters immediately. 7. Judging will be in two parts. First, the written proposals will be assessed by a judging panel comprised of members of AIAA and IGTI organizing committees from the industrial and government communities. Second, the best three teams will be invited to present their work to a second judging panel at a special technical session. The in person presentation will either be at the ASME TurboExpo in Montreal, Canada in June 2015 or the AIAA Propulsion and Energy Forum in Orlando, FL in July The results of the presentations will be combined with the earlier scores from the proposals to determine first, second and third places. 8. Certificates will be presented to the winning design teams for display at their university and a certificate will also be presented to each team member and the faculty/project advisor. The finishing order will be announced immediately following the three presentations. Certificates and recognition in a press release will be the only prizes for this competition. There will be neither prize money nor travel assistance to attend the final presentation.

39 39 II. Copyright All submissions to the competition shall be the original work of the team members. Any submission that does not contain a copyright notice shall become the property of AIAA. A team desiring to maintain copyright ownership may so indicate on the signature page but nevertheless, by submitting a proposal, grants an irrevocable license to AIAA to copy, display, publish, and distribute the work and to use it for all of AIAA s current and future print and electronic uses (e.g. Copyright 20 by. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.). Any submission purporting to limit or deny AIAA licensure (or copyright) will not be eligible for prizes. III. Schedule & Activity Sequences Significant activities, dates, and addresses for submission of proposal and related materials are as follows: A. Letter of Intent November 1, 2014 B. Receipt of Proposal May 1, 2015 C. Proposal evaluations completed - April 30, 2015 D. Round 2 Proposal Presentations & Announcement of Winners June or July See the website for updates as to location of the final presentation. The finished proposal must be received at AIAA Headquarters on or before the date specified above for the Receipt of Proposal (Item B). IV. Proposal Requirements A technical proposal is the most important criterion in the award of a contract. It should be specific and complete. While it is realized that all of the technical factors cannot be included in advance, the following should be included and keyed accordingly: 1. Demonstrate a thorough understanding of the Request for Proposal (RFP) requirements. 2. Describe the proposed technical approaches to comply with each of the requirements specified in the RFP, including phasing of tasks. Legibility, clarity, and completeness of the technical approach are primary factors in evaluation of the proposals. 3. Particular emphasis should be directed at identification of critical, technical problem areas. Descriptions, sketches, drawings, systems analysis, method of attack, and discussions of new techniques should be presented in sufficient detail to permit engineering evaluation of the proposal. Exceptions to proposed technical requirements should be identified and explained.

40 40 4. Include tradeoff studies performed to arrive at the final design. 5. Provide a description of automated design tools used to develop the design. V. Basis for Judging Round 1: Proposal 1. Technical Content (35 points) This concerns the correctness of theory, validity of reasoning used, apparent understanding and grasp of the subject, etc. Are all major factors considered and a reasonably accurate evaluation of these factors presented? 2. Organization and Presentation (20 points) The description of the design as an instrument of communication is a strong factor on judging. Organization of written design, clarity, and inclusion of pertinent information are major factors. 3. Originality (20 points) The design proposal should avoid standard textbook information, and should show independence of thinking or a fresh approach to the project. Does the method and treatment of the problem show imagination? Does the approach show an adaptation or creation of automated design tools? 4. Practical Application and Feasibility (25 points) The proposal should present conclusions or recommendations that are feasible and practical, and not merely lead the evaluators into further difficult or insolvable problems. Round 2: Presentation Each team will have 30 minutes to present a summary of its proposal to the judging panel and answer questions. In addition to the categories above, the presentations will be assessed for clarity, effectiveness and the ability to sell the teams ideas. Scores from the presentation will be added to those from the proposal. The presentation score will be adjusted so that it is worth 30% of the overall value.