Joint AIAA-IGTI Engine Design Competition 2011/12 for Undergraduate Teams

Joint AIAA-IGTI Engine Design Competition 2011/12 for Undergraduate Teams An Engine with High Power Extraction Potential for a Half-Scale Model of a Joint Strike Fighter Request for Proposal August 24, 2011

2 Abstract Designs are requested for an engine to power an unmanned, half-scale model of a Joint Strike Fighter aircraft. The engine should be capable of delivering up to 300 hp of auxiliary power from either the HP or LP spool or some combination throughout the mission, except for take-off. The best modes of operation should be investigated for different segments of the flight. The effects of such high external loads on engine operability should also be examined and remedies suggested to mitigate any adverse behavior. The maximum horsepower that may be taken from the engine without compromising aircraft performance should be established at subsonic & supersonic cruise. The new engine design is to be a scaled version of a generic baseline model, an F100-PW-229, which is supplied. The afterburner temperature will be lower, however. Compromises and limitations of the scaling exercise should be addressed and accounted for whenever possible. GasTurb11 is highly recommended for design & analysis, since this is available at minimum cost to all universities. Design codes developed internally by university departments are acceptable but the use of industrial proprietary codes that are not accessible by all contestants is not permitted. The baseline engine model was generated using an early version of GasTurb11. This code can handle all components except the variable-geometry inlet, and here student teams should resort to engineering fundamentals. A design of industry-quality is not expected and where the tools or the knowledge are not available to address certain issues, assumptions should be made, clearly stated and justified. Primary Contact for RFP Dr. Ian Halliwell Senior Research Scientist, Avetec Inc. Chair - AIAA Gas Turbine Engine Technical Committee Member AIAA Air Breathing Propulsion Systems Integration Technical Committee IGTI Education Committee & IGTI Aircraft Engines Committee e-mail: ianhalliwell@earthlink.net Rules and Guidelines are in Appendix 2

3 CONTENTS Page 1. Introduction 4 2. Design Objectives & Requirements 4 3. Baseline Engine Model 7 3.1 Overall Characteristics 8 3.2 Inlet 17 3.3 Fan 18 3.4 Inter-Compressor Duct 20 3.5 High-Pressure Compressor 20 3.6 Primary Combustor 24 3.7 High-Pressure Turbine 24 3.8 Inter-Turbine Duct 26 3.9 Low-Pressure Turbine 26 3.10 Exhaust System & Augmenter 28 4. Hints & Suggestions 29 5. Competition Expectations 29 References 31 Suggested Reading 32 Appendix 1. Letter of Intent 33 Appendix 2. Rules and Guidelines 34 I. General Rules 34 II. Copyright 35 III. Schedule 35 IV. Proposal Requirements 35 V. Basis for Judging 36

4 1. Introduction Low Bypass Ratio Turbofans - Current & Future Requirements for Military Applications The low bypass ratio, mixed, augmented turbofan engine has dominated military aviation for about four decades. Its mixed abilities include take-off under maximum power, subsonic acceleration & climb under military power, long range subsonic cruise, supersonic penetration of enemy airspace, various combat maneuvers, descent, loiter and landing. In the early part of its history, operational capabilities continued to improve as advances in the constituent technologies - aerodynamics, combustion, materials, controls allowed engine cycles to becomes more complex in the quest to optimize performance at all phases of the mission. More recently, however, restrictive design practices have had to be applied because of noise and emissions regulations, and costs (acquisition, maintenance and operational) have become all important. The ability to augment the thrust by burning additional fuel in the mixed exhaust stream enhances operational performance considerably and avoids the need to size an engine for what may be an infrequent maneuver. While low fuel burn is probably not the most critical characteristic of a military fighter, is is still very important due to the cost & logistics of its transport in wartime. Two factors that are currently of particular interest to the designer of an advanced fighter engine are power extraction and thermal management. As avionics, weaponry and other weaponsplatform systems become more demanding & more sophisticated, considerably more auxiliary power is needed and is taken from the main engine. This is preferable to the design and operation of a separate auxiliary power unit, but it is also especially important that the aircraft propulsion is not compromised. Along with more advanced avionics & weaponry, comes greater heat generation, whose problems are compounded both by operating at higher altitudes & Mach numbers and by the much higher numbers of cyclical, transient and unsteady operations in a mission. The latter generate more heat in the engine bearings and add to the heat load that must be dissipated through the fuel system. In this Request for Proposal, we shall address the effects of high power extraction - its magnitude, its source, and how it influences engine operation. Thermal management will be left for another time! 2. Design Objectives & Requirements An engine design is required for an unmanned half-scale model of a Joint Strike Fighter aircraft. The flight envelope will range from take-off at static sea-level conditions to supersonic cruise at 40,000 feet/mach 1.8. Table 1 shows the minimum net thrust requirements of the aircraft model.

5 Altitude (feet) Mach Number Thrust Demand (lbf) 0 0 9600 0 0.2 8600 5,000 0.4 7600 10,000 0.6 6100 15,000 0.8 4800 20,000 1.0 4400 25,000 1.2 4700 30,000 1.4 5400 35,000 1.6 6200 40,000 1.8 7400 Table 1. Aircraft Minimum Net Thrust Requirements An afterburner will be used where necessary. The maximum value of T 7 is to be 3200 R. (Note that this is less than that used in the baseline engine model! See Section 3.) Full augmentation (max. T 7 ) may be used at take-off but partial after-burning is encouraged whenever possible to achieve the thrust values in Table 1. Throughout the mission an auxiliary power requirement of 67 hp will be needed for general purposes. Throughout the mission (except for take-off), an additional 300 hp of auxiliary power will be required for avionics and weaponry during combat maneuvers, which could occur at both subsonic and supersonic segments. This secondary power offtake may be extracted from either the HP, the LP spool or some combination. A major objective of this exercise is to determine the optimum means of power-off-take in order to maintain thrust to the aircraft and to minimize fuel burn. As the engine operating point changes, any adverse effects of power extraction on engine operation should be noted and mitigating steps should be taken; e.g. fan & compressor stall margins. A second objective is to determine the maximum horsepower that may be taken from the resulting engine without compromising aircraft performance at two mission points - namely (i) Mach 0.9 at 35,000 feet and (ii) Mach 1.4 at 35,000 feet and how this should be split between the two spools. The generic baseline engine model provided should be the starting point for the design exercise. For the current exercise, the engine design point is SLS take-off. The baseline engine model should be scaled, initially by changing the referenced inlet mass flow at the fan face (WR2). Note that although a half-size aircraft is to be powered, this does not

6 necessarily imply a half-size engine! To save weight and cost, the best combinations of engine scale and degree of afterburning that meet the net thrust & power off-take requirements should be sought. Aerodynamic similarity should be maintained, by controlling the referenced rotational speeds (NR1 & NR2) at the engine design point so that the turbomachinery blades and vanes become small replicas of the originals, except for features such as blade clearances and trailing edge thicknesses. In such instances, appropriate adjustments to fan, compressor and turbine efficiencies must be made for Reynolds number effects, where appropriate. Design proposals must include engine mass, engine dimensions, net thrust values, specific fuel consumption, thermal and propulsive efficiencies. Details of the major flowpath components must be given. These include inlet, fan, HP compressor, primary combustor, HP turbine, LP turbine, afterburner, convergent-divergent nozzle, bypass duct, and any inter-connecting ducts. A variable-geometry, two-dimensional inlet must be designed that allows the engine to operate satisfactorily over the mission envelope. It is expected that the inlet will incorporate external oblique shocks. Dimensions should be provided in either a table or in a figure at take-off and max cruise conditions. An estimate of inlet mass is also required, so appropriate inlet materials and thicknesses must also be selected. A variable-geometry, annular exhaust nozzle must be designed. The documentation requirements are the same as for the inlet. Figure 1. Flight Envelope for the Half-Scale Baseline Engine Model

7 Figure 1 shows an example of the Flight Envelope for the half-scale baseline engine, with the HP spool speed fixed at 100%. Full after-burning has been set (T 7 = 3200 R). A constant power off-take of 67 hp has been applied. The minimum equivalent air speed is 350 knots, the maximum equivalent air speed is 750 knots. Contours of net thrust have been plotted and these are intended to serve as a rough guide to what may be achieved. 3. Baseline Engine Model The baseline engine is a generic model of a F100-PW-229, generated from publically-available information (Reference 1). The real engine is a low bypass ratio, augmented turbofan, used on different versions of the F-15 and F-16 fighter aircraft with cruise capabilities in excess of Mach 2. The data is released courtesy of Avetec Inc. (It should be noted that this engine model was generated from an early beta version of the code and some variations too the input may now be needed!) Figure 2 contains a schematic of a low bypass ratio, mixed, augmented turbofan engine with station numbers. Figure 2. A Mixed, Low Bypass Ratio, Augmented Turbofan Engine with Calculation Stations & Cooling Flows

8 3.1 Overall Characteristics Table 1 contains a summary of basic engine characteristics, taken directly from Reference 1. Design Feature F100-PW-229 Engine Type Augmented, turbofan Number of fan/compressor stages 3, 10 Number of HP/LP turbine stages 2, 2 Combustor type Annular Maximum net thrust at sea level 29,000 lbf Specific fuel consumption at max. power 1.94 lbm/hr/lbf Overall pressure ratio at max. power 32.0 Max. envelope diameter 46.2 inches Max. envelope length 191.2 inches Dry weight less tail-pipe 3,395 lbm Current Applications F-15E/I/S, F-16C/D Table 1: Baseline Engine: Basic Data, Overall Geometry & Performance Major Design Parameters In a regular turbofan engine, the primary four 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). Even in a low bypass ratio system, we differentiate between the fan pressure ratios in the core & bypass streams. In a mixed augmented turbofan engine, the maximum temperature in the afterburner (T 7 in Figure 1) must also be selected. Table 2 is the Basic Input for the GasTurb11 model of the F100-PW-229 baseline engine. It also provides some guidelines for the new engine design.

9 Table 2: Basic Input Of the five primary design variables, only the overall pressure ratio is given in Table 1. 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 static sea level take-off conditions as targets. Of course, by definition, this operating condition also corresponds to the engine design point. Table 2 above contains most of the primary input parameters for the engine cycle. Some of the secondary inputs are also discussed here while the rest are covered below. The upper portion of Table 2 defines the ambient conditions upstream of the engine inlet (i.e. at Plane 1). The first row of the second block of Table 2 allows for a 1% total pressure loss between the inlet leading edge and the fan face. The inner and outer fan pressure ratios are then selected separately; since there is more blade speed at the fan tip than at its hub, the inner & outer fan pressure ratios have

10 been set at 2.5 & 3.0. A 1% total pressure loss is then accounted for in the duct between the fan and the HP compressor. Knowing that the required overall pressure ratio is 32.0 results in a pressure ratio across the rear block of 32/2.5 = 12.8 and this also is an input value. (The intercompressor duct pressure loss (0.99) is now accounted for so the rear block pressure ratio is now given by 32/2.5/0.99 = 12.933). Next a 3% total pressure loss is assumed in the bypass duct, followed by an inter-turbine duct loss of 2%. The fan pressure ratios at inner & outer strata were accommodated easily by a three-stage fan design. The 12.8 pressure ratio on the rear block also corresponded adequately to the ten-stage HP compressor mandated in Table 1. Continuing with the input description, the design bypass ratio was set initially at 1.0. Subsequently the cycle calculation was run in optimize mode, where minimum specific fuel consumption was sought while varying the bypass ratio. This resulted in the final bypass ratio value of 0.930 seen in Table 2. A value of 2880 R for the turbine entry temperature was guessed as being reasonable for an advanced military engine of this vintage with limited cooling capacity and an expected long life for the HP turbine (say 5,000 hours). 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 User Guide of Reference 2. Without expert knowledge, this is best left alone! The remaining parameters in Table 2 may be considered as secondary influences and are discussed briefly below. The final major input parameter, the augmenter temperature (T 7 ) is not given in Table 2 but is inserted via the reheat tab on the same screen. A value of 3600 R has been used in this engine model. Clearly this is significantly higher than the turbine entry temperature emanating from the primary burner. However, the reason that this is perfectly viable is that there are no rotating parts downstream. As long as the integrity of the augmenter liner can be guaranteed by either the use of an adequate cooling air supply or suitable materials, then very high afterburner temperatures are sustainable. In addition to a T 7 value of 3600 R, the reheat design efficiency of 90% was selected, along with a quantity of nozzle cooling air of 5% W 16 and a reheat part load constant of 0.5. Note that even though augmentation is very inefficient, it is preferable to scaling up the whole engine for mission segments of limited duration. 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 2 (although this is in fact zero), however the secondary flows indicated in Figure 1 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 are also specified in Table 2 by inserting the appropriate pressure ratios across the inter-compressor duct, the interturbine 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

11 specifically listed in Table 2, but may be reviewed in the output summary presented later in Table 3. 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 Table 3. Another option is available that allows GasTurb11 to estimate turbine efficiencies from data supplied; this leads to values of stage loading and flow coefficients, which are then used in a Smith Chart (Reference 3), assuming an equal work spilt between stages. It is recommended that this be used in this exercise. Power Off-take: All engines have power extracted - usually from the HP spool via a tower shaft that passes through an enlarged vane in the main frame. 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 military application currently under consideration, a great deal of auxiliary power is needed for weaponry and avionics and this usage is growing rapidly in modern aircraft. We have selected a power off-take of 67 hp from our engine. This value should be retained throughout the current exercise, in addition to the 300 horsepower off-take requested for combat purposes. Mixer Efficiency: Mixer efficiency quantifies the degree of mixing that is achieved at plane 163 between the core flow and the bypass flow. It can be shown analytically that thrust is maximized if the mixing is total. In order to do this a large & heavy active mixer would be required; therefore an appropriate compromise is arrived at, since a large mixer means a heavier engine that requires more thrust an uphill spiral! In order to optimize whatever mixing is mechanically possible, the designer must also ensure that the (static) pressures are equalized in the flows leaving the engine core and bypass duct by trading the work balance between the high- and low-speed spools and adjusting annulus areas to effect velocities. The bypass ratio plays a key role here. 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 GasTurb 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. We define a rubber engine initially where performance is delivered in terms of a net thrust of 29,000 lbf given in Table 1 once the engine scale has been determined. We also have a dimensional envelope to fit into, namely a maximum casing diameter of 46.2 inches and a maximum length of 191.2 inches. 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 0.55 was found to be appropriate. This sets the general radial dimension for the complete engine, although in fact downstream of the fan, the entry radius of the compressor is also determined by an input radius

12 ratio. 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 s probably OK!), with final manipulations being added as the target dimension is approached. Materials & Weights As far as possible, use was made of the materials database in the GasTurb11 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-carbon products, 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 F-15/F-16 applications, where reliability is critical. However, within the component models, material 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 4, interpolating and extrapolating where necessary. In GasTurb11 component weights are calculated by multiplying the effective volumes by the corresponding material densities. Of course, only the major elements which are directly designed 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 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 GasTurb manual, but I reduced this to a net mass factor of 1.1 in Table 4, mainly because it got me closer to the gross engine weight I was looking for!

13 Table 3 is the Output Summary Table from Gasturb11 for the F100-PW-229 baseline engine. Table 3: Baseline Engine Output Summary Table 4 is a more detailed overall output table from Gasturb11 for the F100-PW-229 baseline engine. Figures 2 & 3 contain a cutaway of the baseline engine and a plot of the GasTurb11 model respectively.

Table 4: Baseline Engine Detailed Output 14

Figure 2. The F100-PW-229 Baseline Engine 15

Radius [in] 16 Figure 3. GasTurb11 Model of the Baseline Engine

17 Some details of the component models now follow. These are intended to help with its duplication as well as with the design of the new scaled engine. 3.2 Inlet It should be noted that the inlet described here is not the variable-ramp device that interfaces between the atmospheric conditions and the location of the throat, as requested in the penultimate bullet of Section 2, but is merely the structural duct downstream of normal shock location in supersonic operations. The inlet is designed with a conic centerbody and a set of struts (see Figure 3). The distinction between a strut assembly in the inlet, which also functions as a front frame, and a set of inlet guide vanes contained in the fan model needs to be made mainly in regard to how the GasTurb inlet model functions. Both approaches were tried but it was found that the use of a fan IGV did not permit the conic centerbody to be constructed correctly because the IGV inner wall was at a constant radius. The use of an upstream strut or frame allowed a linear closure of the inlet bullet at the engine centerline without an elliptic profile being incorporated. In engineering practice, the presence of a three-stage fan foretold the need for a front frame. This is because a threestage fan cannot be cantilevered from a bearing supported by the engine mainframe, located somewhere aft between the fan and HP compressor, but must be centered via its own bearing; hence the need for a front frame. The outer diameter of the inlet has been determined from that of the fan. Table 5. Inlet Design Pertinent characteristics of the inlet are shown in Table 6. At 133 lbm, 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. This should accommodate the dynamic heating effects of high-speed operation. It is noteworthy that the GasTurb inlet is merely the portion of the casing (plus centerbody) immediately upstream of the fan. In practice, a variable-geometry inlet (part of the engine/airframe integration) is necessary for the range of Mach numbers encountered in an F- 15/16 mission. This permits the necessary flow diffusion from supersonic ambient conditions down to a subsonic value at the fan face and this is accomplished via a series of oblique shocks

18 followed by a final normal shock at the inlet throat. Of course this is made up of another significantly large structure of considerable complexity and mass, whose configuration can be changed in order to optimize total pressure recovery throughout the mission. However, the GasTurb model begins at the upstream flange and the real inlet must be accounted for elsewhere as part of the airframe which fortunately lets us off the hook! 3.3 Fan Table 6. Fan: Detailed Overview Note that the value of the tip speed at the top of Table 5, in conjunction with the fan tip radius, that sets the rotational speed of the LP spool. Table 7. Fan General Output

Table 8. Fan Disks 19

20 3.4 Inter-Compressor Duct Table 9. Inter-Compressor Duct 3.5 High-Pressure Compressor Table 10. High Pressure Compressor: Detailed Overview Again, we effectively set the speed of the HP spool via the tip speed and associated radius!

21 Table 11. High Pressure Compressor: General Output Table 12. High Pressure Compressor: Annulus

Table 13a. High Pressure Compressor: Stages 1 6 Disks (A bug in the old Beta version of the code has resulted in negative temperatures in the disks) 22

Table 13b. High Pressure Compressor: Stages 5-10 Disks (A bug in the old Beta version of the code has resulted in negative temperatures in the disks) 23

24 3.6 Primary Combustor Table 14. Combustor 3.7 High-Pressure Turbine Table 15. High Pressure Turbine: General Output

Table 16. High Pressure Turbine: Disks (The negative stress margins in stage 2 were subsequently corrected) 25

26 3.8 Inter-Turbine Duct Table 17. Inter-Turbine Duct 3.9 Low-Pressure Turbine Table 18. Low Pressure Turbine: General Output

Table 19. Low Pressure Turbine: Disks 27

28 3.10 Exhaust System & Augmenter The cone ends in the exhaust duct Table 20. Core Exhaust Table 21. Augmenter Table 22. Convergent-Divergent Nozzle

29 4. Hints & Suggestions You should first model the baseline engine with the same software that you will use for the new engine design. Your results may not match the baseline engine model exactly but will provide a valid comparison of weights and performance for the new concept engine. The efficiencies of the turbomachinery components may be assumed to be the same as those of the baseline engine, and input directly or the calculate efficiency mode GasTurb11 may be invoked. Instruction in the use of GasTurb11 may be arranged. As stated in the Abstract, the use of design codes from industrial contacts, that are not accessible to all competitors, is not allowed. Even though the date for submission of Letters of Intent is stated as November 1, 2011 on pages 33 and 35, it is recommended that teams who know that they will enter the competition inform either AIAA or Dr. Ian Halliwell (ianhalliwell@earthlink.net) as soon as possible, so that assistance may be given or access to design codes may be arranged. 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. The responses will be made available to all participants in the competition via an AIAA website. 5. Competition Expectations The existing rules and guidelines for the AIAA Foundation Student Design Competition should be observed and these are supplied in Appendix 2. In addition, the following specific suggestions are offered for the Engine Design Competition. 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:

30 Definition and justification of critical mission point(s) that drive the candidate propulsion system design(s). 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 flowpath that shows how the major components fit together, with emphasis on operability at different mission points. A clear demonstration of design feasibility, with attention having been paid to technology limits. Stage counts. 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 flowpath 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 also 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.

31 References 1. Aerospace Source Book. Aviation Week & Space Technology. January 15, 2007. 2. GasTurb 11: A Design & Off-Design Performance Program for Gas Turbines <http://www.gasturb.de> Joachim Kurzke, 2007. 3. A Simple Correlation of Turbine Efficiency S. F. Smith Journal of the Royal Aeronautical Society. Volume 69. 1965. 4. Aeronautical Vest Pocket Handbook. Pratt & Whitney Aircraft. Circa 1980

32 Suggested Reading 1. Gas Turbine Theory H.I.H Saravanamuttoo, G.F.C Rogers &.H. Cohen, Prentice Hall. 5 th Edition 2001. 2. Aircraft Engine Design J.D.Mattingly, W.H. Heiser, & D.H. Daley AIAA Education Series. 1987. 3. Elements of Propulsion Gas Turbines and Rockets J.D. Mattingly. AIAA Education Series. 2006. 4. Jet Propulsion N. Cumpsty. Cambridge University Press. 2000. 5. Gas Turbine Performance P. Walsh & P. Fletcher. Blackwell/ASME Press. 2 nd Edition, 2004. 6. Fundamentals of Jet Propulsion with Applications Ronald D. Flack Cambridge University Press. 2005. 7. The Jet Engine Rolls-Royce plc. 2005.

33 Appendix 1. Letter of Intent 2011/20012 Joint AIAA IGTI Undergraduate Team Engine Design Competition Request for Proposal: An Engine with High Power Extraction Potential for a Half-Scale Model of a Joint Strike Fighter Title of Design Proposal: Name of School: Designer s Name AIAA or ASME Graduation Date Degree ID or Member # Team Leader Team Leader E-mail In order to be eligible for the 2011/2012 Joint AIAA-IGTI Undergraduate Team Engine Design Competition, you must complete this form and return it to the AIAA Student Programs Coordinator before November 1, 2011, at AIAA Headquarters, along with a one-page Letter of Intent, as noted in Appendix 2, Section III, Schedule and Activity Sequences. For any nonmember listed above, a student member application and member dues payment to AIAA should also be included with this form or submitted to ASME, with a note attached. Signature of Faculty Advisor Signature of Project Advisor Date Faculty Advisor Printed Project Advisor Printed Date

34 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 AIAA or ASME/IGTI branch or at-large Student Members per entry. 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. 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 and signatures. 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 the 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 at the 48 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference in Atlanta in July 2012. Airfare and lodging expenses will be partially covered for the invited teams and their advisors. The results of the presentations will be combined with the earlier scores to determine first, second and third places. 8. The prizes shall be: First place-$2,500; Second place-$1,500; Third place-$1,000 (US dollars). 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.

35 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, 2011 B. Receipt of Proposal April 1, 2012 C. Proposal evaluations completed - April 30, 2012 D. Proposal presentations & Announcement of Winners at JPC in Atlanta, GA July 2012 Teams intending to submit a proposal must submit a one page Letter of Intent along with the signed attached Intent Form (Item A) on or before the date specified above, to the following address: AIAA Student Programs 1801 Alexander Bell Drive Suite 500 Reston, VA 20191-4344 The CD containing the finished proposal must be received at the same address on or before the date specified above for the Receipt of Proposal (Item B). IV. Proposal Requirements The technical proposal is the most important factor 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:

36 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. 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 A. 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. B. Presentation Each team will have 40 minutes to present a summary of its proposal to the judging panel. In addition to the categories above, the presentations will be assessed for clarity, effectiveness and the ability to sell the teams ideas.