Joseph M. Roche, Donald T. Palac, James E. Hunter, David E. Myers, and Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Daniel N. Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

David R. McCurdy and Kevin T. Dougherty

QSS Group, Inc., Cleveland, Ohio

Investigation of Exoskeletal Engine Propulsion

System Concept

NASA/TM—2005-213369

August 2005
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Joseph M. Roche, Donald T. Palac, James E. Hunter, David E. Myers, and Christopher A. Snyder

Glenn Research Center, Cleveland, Ohio

Daniel N. Kosareo

ZIN Technologies, Inc., Brook Park, Ohio

David R. McCurdy and Kevin T. Dougherty

QSS Group, Inc., Cleveland, Ohio

Investigation of Exoskeletal Engine Propulsion

System Concept

NASA/TM—2005-213369

August 2005

National Aeronautics and

Space Administration

Glenn Research Center
Available from

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This work was sponsored by the Low Emissions Alternative

Power Project of the Vehicle Systems Program at the

NASA Glenn Research Center.

Acknowledgments

The authors wish to acknowledge the assistance of those who aided in the preparation of this document. Christos C.

Chamis, Christopher Dellacorte, and Andrew J. Provenza of NASA Glenn Research Center provided information

on technologies and systems that constitute the exoskeletal engine concept. Louis M. Larosiliere provided

information on turbomachinery aerodynamics and its application to the exoskeletal design. David

Cherry, Eric Estill, Ken Willgoose, and Rob Beacock of General Electric Aircraft Engines mined

the archives for data on a concept with some precedence for an exoskeletal design, the

unducted fan. Ian Halliwell of Modern Technologies Corporation provided background

and the results of previous exoskeletal engine preliminary design studies.

Reza N. Zinolabedini provided NASTRAN eigenvalue buckling solutions

for the stationery core structure under compressive thrust loading.
NASA/TM—2005-213369 1 

Investigation of Exoskeletal Engine Propulsion System Concept 

Joseph M. Roche, Donald T. Palac, James E. Hunter, David E. Myers, and Christopher A. Snyder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Daniel N. Kosareo 

ZIN Technologies, Inc. 

Brook Park, Ohio 44142 

David R. McCurdy and Kevin T. Dougherty 

QSS Group, Inc. 

Cleveland, Ohio 44135 

Summary

 An innovative approach to aircraft gas turbine engine design that was disclosed in 1997 involves 

mounting compressor and turbine blades to an outer rotating shell. Designated the exoskeletal engine, the 

concept results in compression (especially preferable to tension for high-temperature ceramic materials, 

generally) as the dominant compressor and turbine blade force. 

 It was determined that the feasibility of an exoskeletal engine lay in the challenges of structural and 

mechanical design (as opposed to cycle or aerothermodynamic design), so the focus of this study was the 

development of a structural-mechanical definition of an exoskeletal concept, using the Rolls-Royce 

AE3007 regional airliner all-axial turbofan as a baseline. The effort was further limited to the definition 

of an exoskeletal high-pressure spool concept, where the major structural and thermal challenges are 

represented. A finite-element model of an exoskeletal engine high-pressure spool shell and bearing 

support structure was developed. Loads and deflections were calculated and were compared with 

material strengths at the temperatures expected for each portion of the shells. Magnetic and foil bearing 

system concepts were also defined. The mass of the high-pressure spool was calculated and compared 

with the mass of the comparable components of the AE3007 engine. It was found that the exoskeletal 

engine rotating components have the potential for significant weight savings over the rotating 

components of conventional engines. However, bearing technology development is required for this mass 

savings to emerge at an engine system level, since the mass of existing bearing systems would exceed 

rotating machinery mass savings. It is recommended that once bearing technology is sufficiently 

advanced, a “clean sheet” preliminary design of an exoskeletal system be accomplished. This design 

would take best advantage of exoskeletal engine features from the structural, mechanical, aerodynamic, 

thermal, and engine cycle perspectives. This will better quantify the potential for the exoskeletal concept 

to deliver benefits in mass, structural efficiency, and cycle design flexibility. 

Introduction 

 To improve conventional turbine and compressor designs, it is hypothesized that ceramic materials 

would provide greater operating efficiency through higher operating temperatures and lighter engine 

weight. Some problems with ceramic materials are lower tensile strength and durability and damage 

tolerance when compared with refractory metals and nickel-based alloys, such as titanium, HastelloyTM

and WaspaloyTM that are currently used in conventional engines. Ceramics behave well in compressive 

loading situations where brittle fracture is minimized. A novel turbine design approach that relies on this 

characteristic of ceramic composites is the exoskeletal engine (Christos C. Chamis, August 9, 2002, 
NASA/TM—2005-213369 2 

NASA Glenn Research Center, Cleveland, OH, personal communication), which was disclosed by 

Christos Chamis of NASA Glenn Research Center (GRC) in 1997. The exoskeletal concept exchanges 

the conventional internal rotating shaft with a stator and the stator with a rotating drum. Essentially, this 

concept turns a conventional gas turbine engine inside out. A cross-sectional diagram of an early 

conceptual exoskeletal engine configuration is shown in figure 1. 

 The exoskeletal engine does not in principle represent an engine cycle change, although cycle 

parameters could be affected somewhat in implementation. Rather, the exoskeletal concept deals 

primarily with engine structural and mechanical design. Some of the characteristic differences from 

conventional engine technology are as follows: 

 (1) Compressor and turbine blades are mounted to the inside of a drum rotor. Stators and combustors 

are mounted to a stationary hub. Blades are in compression instead of tension because of rotational 

inertia. Torque is transferred from turbines to compressors through the rotating drums. 

 (2) The drum rotor is supported by bearings between the outside of the rotor and an outer shell. In a 

multispool engine, the high-pressure drum rotor bearings may ride directly on the low-pressure spool 

drum rotor. In a partially exoskeletal engine, thrust and other forces are transferred through bearings 

between drum rotors and shells or hubs. 

 (3) For engines of sufficient size, a central core space (inside the stator rings) may be open 

(see fig. 2). 

 The motivations for considering an exoskeletal engine have included the opportunity to 

 (1) Reduce the part count and engine weight, achieve higher engine operating temperatures (or 

reduce/eliminate cooling), and increase component life through use of lightweight composite materials 

(particularly ceramic blades in compression and rotating wound shells under hoop stresses) in 

applications that are aligned with their strengths 

 (2) Reduce engine noise through management of the resulting inverted velocity profile (central flow 

velocity lower than surrounding flow velocity) 

 (3) Insert another engine cycle, such as a ramjet, into the open core space 
NASA/TM—2005-213369 3 

Related Prior Efforts 

 General Electric unducted fan.⎯The General Electric unducted fan (UDF), developed as part of 

NASA’s advanced turboprop program in the 1980s, represents a successful, if not entirely similar, flightproven predecessor to the exoskeletal engine concept as shown in figure 3. The UDF engine consisted of 

a standard high-pressure engine spool to which a low-pressure turbine (LPT) was added to directly drive 

a set of counterrotating highly swept propellers, also called unducted fan blades. The aft row of fan 

blades was driven by a conventional axial turbine, but the forward row of fan blades was driven by 

several rows of turbine blades mounted at their outer edge to a rotating drum. On the UDF, this drum was 

connected to and supported by a more conventional disk (ref. 1). Because the UDF low-pressure turbine 

drove a set of high-speed propellers (or a low-speed unducted fan), wheel speeds were low compared 

with turbofan engines. Centrifugal forces were not a factor in blade or drum design (David Cherry, 

Dec. 5, 2002, General Electric Aircraft Engines, Cincinnati, OH, private communication). In this respect, 

the UDF represents a significantly different set of design objectives and experience from those that 

would be relevant to an exoskeletal engine concept. However, the UDF does represent the successful 

implementation in flight of some elements of an exoskeletal engine architecture. 

 Modern Technologies Corporation exoskeletal engine concept studies.⎯NASA-sponsored studies 

of the exoskeletal engine concept were conducted by Modern Technologies Corporation between 1999 

and 2001. These studies examined exoskeletal design concepts for 2000-, 5000-, and 25 000-lb thrustclass engines. Preliminary design concepts consisted of cycle definition and stage number and size 
NASA/TM—2005-213369 4 

definitions (no structural analyses were performed). The studies pointed out the critical need for 

exoskeletal engine preliminary design guidelines and structural constraints on preliminary design because 

of the departure from conventional engine design practices (ref. 2). 

 Honeywell International, Inc., exoskeletal engine study.⎯In early 2000, Honeywell International, 

Inc., developed an exoskeletal engine concept based on an existing engine (NASA Exoskeletal Engine 

Study Final Report, Honeywell International, Inc., May 23, 2000). Honeywell selected the AS900 engine 

(approx. 6500-lb thrust at sea level static conditions) as a basis for comparison. A single-spool concept 

was assembled but was rejected because of performance, operability, and engine-starting limitations as 

compared with a two-spool configuration. A two-spool engine concept was assembled. A layout of this 

concept was created, and an engine cycle simulation was assembled sufficient to calculate the 

performance of engine components. Hoop stresses in the drum rotors were calculated and were used to 

identify material options and weights. High-pressure spool (HPS) rotation speeds necessary to produce 

performance similar to the AS900 required bearings operating at approximately 9 million mm-rpm, 

beyond the state of the art for bearings. The performance of a lower speed high-pressure spool was 

calculated for a rotation speed within state-of-the-art bearing capabilities of 6 million mm-rpm. Engine 

weights were calculated for rolling-element and magnetic bearing options, and weights and performance 

were compared with the AS900 engine. Major conclusions of this effort were 

 (1) An exoskeletal engine concept that delivers similar performance to an existing conventional 

engine is feasible but would have a radius approximately 30 percent smaller and would be approximately 

50 percent longer. 

 (2) Exoskeletal rotating machinery is lighter (in concept) than that for conventional engines. 

 (3) Bearing system weight for an exoskeletal engine increases the total weight of the engine to 20 to 

25 percent greater than that of a conventional engine. 

 It should be noted that the Honeywell AS900 engine has a centrifugal high-pressure compressor 

(HPC). Centrifugal compressors generally add weight and diameter but decrease the length of gas turbine 

engines. This creates some ambiguity about the directness of the comparison between the AS900 engine 

and a completely axial exoskeletal engine. However, the Honeywell study represented the first known 

assessment of the exoskeletal concept applied to an engine as a system. The findings of the Honeywell 

study indicated the importance of engine systems approach in any subsequent investigations, as well as 

the predominance of engine structures, materials, and bearings in the application of the exoskeletal 

concept. 
NASA/TM—2005-213369 5 

Scope and Rationale 

 The present investigation focused on development and analysis of a gas turbine engine high-pressure 

spool of an exoskeletal engine concept. Goals of this investigation were to 

 (1) Determine the weight of an exoskeletal engine HPS concept. The HPS, rather than an entire 

engine system, was selected for study because 

 (a) The HPS typically represents the greatest challenges to structural, thermal, and bearing design 

in an engine, particularly for the exoskeletal concept characteristics of interest (e.g., high rotation speeds 

in compressor and turbine and high turbine temperatures). A preliminary calculation of hoop stresses in 

shells identified that small- to medium-size engine high-pressure spools, operating at comparable rotation 

speeds with existing engines, would challenge limits of existing structures and materials. A medium-class 

(approx. 10 000-lb thrust) engine high-pressure spool was chosen as the design example for this 

investigation. 

 (b) The analysis of an HPS is manageable in a reasonable study period, yet is enough of an engine 

system in itself to provide system-level answers. 

 (2) Perform requisite analysis to ensure that the concept is realistic yet aggressive. Analysis is needed 

to establish feasibility and traceability of weights of an exoskeletal HPS concept components. However, 

existing gas turbines benefit from 40 years of design refinement, so an aggressive approach gives the 

benefit of the doubt to an exoskeletal approach appropriate to making a comparison with an existing 

engine HPS. 

 (3) Compare the weight of an exoskeletal HPS concept with a representative existing gas turbine 

engine HPS. 

 (4) Identify technology challenges of the exoskeletal concept. 

 Several aspects of the exoskeletal engine were considered beyond the scope of this investigation. 

This investigation did not attempt to quantify the benefits of an exoskeletal engine for noise reduction or 

for compatibility with other engine cycles, such as a ramjet. It also did not include an optimization of an 

exoskeletal design, either structurally, aerodynamically, or thermodynamically, for a specific engine class 

or set of requirements. It is expected that these issues are more appropriately dealt with once an approach 

to exoskeletal engine structural and mechanical systems design and resolution of the inherent technology 

challenges have been accomplished.

 Appendix A provides the output of the Weight Analysis Turbine Engine (WATE) computer program. 

The WATE code estimates the total engine weight by using the weights of the various components and/or 

subcomponents, based upon user inputs and conditions experienced. Appendix B presents the output of 

the NASA Engine Performance Program (NEPP), which develops a one-dimensional, steady-state 

thermodynamic cycle model and generates compressor and turbine maps. Appendix C presents the input 

data for an ANSYS structural analysis of the exoskeletal rotor when it is spinning at the over-speed 

condition of 16 400 rpm. This input file contains preprocessing, solution, and postprocessing commands 

to create the model, perform the analysis, and plot the results. Appendix D contains the input data for the 

Pro/Engineer® models of the compressor and turbine rotor blades. The input data for the exoskeletal 

compressor rotor blade shapes are based on the NACA 65A010 compressor airfoil shape. 

Symbols

AR blade aspect ratio 

c chord 

DN diameter-number of revolutions (for bearings), millions of mm- rpm 

E modulus of elasticity, lb/in.2
NASA/TM—2005-213369 6 

F force, lbf

G shear modulus of elasticity, lb/in.2

g gravitational constant 

H blade height, in. 

I area moment of inertia, in.4

J polar moment of inertia 

K matrix stiffness, lb/in. 

k effective stiffness, lb/in. 

L stage length, in. 

M matrix mass, lbm

m mass , lbm

N rotational speed, rpm

NB number of blades 

PR pressure ratio 

r radius, in. 

V shear force, lbf

W weight, lbm

WB blade weight (also includes fir tree), lbm

x axial location 

α coefficient of thermal expansion (CTE) 

γ weight density, lb/in.3

δ radial deflection, in. 

θ angular deflection, rad 

θ stagger angle, deg 

μ Poisson’s ratio 

ρ material density, lbm/in.3 

σ stress, lb/in.2

φ camber line angle, deg 

ω rotational speed, rad/sec 

Subscripts: 

ACC accessories weight (also includes other weights that show up but are difficult to assign to one 

 particular part) 

a inner 

B blade 

b outer 

c corrected 

cu ultimate compressive allowable 

cy ultimate compressive yield allowable 

disk disk 

e exit 

H hub 

i inlet 
NASA/TM—2005-213369 7 

lc maximum allowable compressive stress in fiber direction 

1t maximum allowable tensile stress in fiber direction 

max maximum 

min minimum 

N nuts and bolts (also used to bolt disks together) 

n natural or local 

RD rotor drum 

r radial 

row row 

stage stage 

su ultimate shear allowable 

T tip 

t tangential 

tu ultimate tensile allowable 

ty tensile yield 

vm Von Mises 

τ shear 

1 maximum principal stress, klb/in.2

2 minimum principal stress, klb/in.2

Assumptions, Key Challenges, and Approach 

Assumptions 

 The first assumption was that the application of conventional aerothermodynamics design principles 

to an exoskeletal engine architecture was sufficient to identify the limiting structural and materials 

challenges. As mentioned above and in reference 2 (Halliwell), the design of exoskeletal engine blades 

and flow paths to deliver performance on a par with highly optimized conventional engines is likely to 

require significant study and development of design guidelines. For example, exoskeletal rotating blade 

tips are at a smaller radius than the blade root, which results in tip leakage at the engine hub increasing 

with rotational speed. Rotating blades supported at the outer radius and stators supported at their inner 

radius necessitate a structural design opposite that of conventional engines, affecting thickness, 

curvature, and blade dynamics. In an exoskeletal engine, the outer wall of the engine rotates with the 

blades, representing a large windage area (compared with a conventional shaft). The impacts on engine 

performance of these and other changes from conventional engine design are unknown. However, the 

exoskeletal engine concept proposes no intended deviations from the aerothermodynamic cycle design 

heritage of conventional engines. In lieu of design details, an earlier examination of the feasibility of the 

exoskeletal engine concept found it acceptable to transfer engine cycle parameters directly from a 

baseline conventional engine (Exoskeletal Technology Evaluation for Use in Aeropropulsion 

Applications. Presented to the Ultra-Efficient Engine Technology Program, Feb. 28, 2001). It was 

therefore assumed that any variations in engine aerothermodynamic cycle design are of small order 

compared with the issues of viability presented by structural-mechanical design and that an exoskeletal 

engine design can be assumed aerothermodynamically identical to a conventional engine for the purposes 

of this investigation. 

 The second assumption was that turbine engine subsystem and components currently in use can be 

applied to an exoskeletal engine at their existing performance levels. Typical gas turbine engine 

components, such as controls, actuators, and combustors, may pose significant engineering challenges in 
NASA/TM—2005-213369 8 

their adaptation to an exoskeletal engine. For this investigation, however, the expectation was that no 

new technology would be required to apply these subsystems and components to an exoskeletal engine. 

Key Challenges 

 Development of structural and mechanical exoskeletal HPS concept.⎯The central idea of an 

exoskeletal engine concept involves a change in the structural and mechanical design of the rotating 

machinery in a gas turbine engine. The concept is not proposed as a change to the engine cycle or 

aerodynamics, other than that necessary to accommodate changes to the structural and mechanical 

design. Thus, the system-level feasibility questions about the exoskeletal engine concept lie with the 

structural and mechanical architecture. The outer rotating shell of an exoskeletal system must withstand 

stresses due to the inertia of rotation, contain engine pressure loads, react to blade loads, and 

accommodate the engine nominal and contingency environments. It must also accommodate engine 

assembly and maintenance. 

 Development of system-integrated bearing concept for exoskeletal HPS.⎯The exoskeletal 

approach inherently involves large portions of an engine system rotating at larger radii than those in a 

conventional engine. The development of an appropriate bearing concept is a challenge that involves 

 (1) Integration with the system structural concept to minimize bearing DN

 (2) Consideration of existing and emerging bearing technologies, including noncontact bearings 

 (3) Consideration of both nominal and off-nominal engine conditions, including startup, 

maneuvering, and landing 

Approach 

 A conceptual analysis was deemed sufficient to determine exoskeletal structural-mechanical viability 

and to compare the weight of a conventional engine with its exoskeletal counterpart. This approach 

included both identifying the key characteristics driving the viability and weight and selecting the 

pertinent engine subsystems for design and analysis. A representative conventional engine was selected 

and the exoskeletal analogue was created. This approach minimized the design and analysis effort while 

still providing the understanding of the engineering issues involved with the exoskeletal architecture. In 

detail, the process can be described as follows: 

 Identification of suitable conventional gas turbine engine for comparison with exoskeletal 

concept.⎯A valid comparison of an exoskeletal approach and the conventional approach required 

identification of an existing gas turbine engine with the following characteristics: 

 (1) Representative of state-of-the-art gas turbine engine design, especially high-pressure spool 

rotation speeds and temperatures 

 (2) Availability of design information, allowing modeling and adaptation of structures and 

mechanisms to an exoskeletal approach 

 (3) Fully axial design (no centrifugal compressors or turbines) 

Design of exoskeletal analogue HPS concept.⎯A conceptual design of an exoskeletal engine highpressure spool was developed based on the NASA model of an existing engine. The conceptual design is 

described in the section Description of Exoskeletal System Concept. The approach to this design had 

neither a precedent nor design guidelines for exoskeletal components. Materials and structures concepts 

were defined consistent with the analogous environments from the conventional HPS example, consistent 

with an aggressive but realistic concept philosophy. 

Calculation of weight of exoskeletal analogue HPS concept.⎯The weight of the exoskeletal HPS 

was calculated based on the design concept, as were the weights of shells and the stationary core. Blade 

weights were based on conventional engine blade weights but were modified for the exoskeletal 
NASA/TM—2005-213369 9 

geometry. The weights of components, such as combustors, that were largely unaffected by the change to 

an exoskeletal design were adopted from existing engines. 

Verification of exoskeletal HPS concept structural integrity.⎯Structural analyses were performed 

to verify the structural integrity of critical exoskeletal high-pressure spool components, such as drum 

rotors, bladed rings, and bearing supports. The intent was to show that the weight estimates were 

aggressive but reasonable. Finite-element models were not of sufficient detail to perform complete 

structural analyses considering the limited thermal analysis that was performed. The operating thermal 

environment was, however, considered in the selection of materials and in choosing an appropriate 

allowable stress range for applied mechanical loads. This approach was considered adequate for the 

purposes of this investigation. High-fidelity modeling and refinement of the environment to include all 

aerodynamic, inertial, and thermal loads to a greater degree of accuracy would be required prior to 

commitment to an exoskeletal engine development.

 Identification of technology challenges.⎯This investigation attempted to discover not just whether 

an exoskeletal approach to gas turbine engine design is viable but also the critical characteristics upon 

which viability depends. Where new or improved technologies were important to the viability of the 

exoskeletal approach, they were identified. A preliminary assessment was made of the potential of new 

technologies to satisfy the requirements of the exoskeletal approach. 

Architecture Description 

 Since the exoskeletal concept presents an aggressive departure from conventional turbine engines, 

it is imperative to look at the concept using a systems approach. A solid model of the concept was 

generated as the first step of the analysis and weight determination of the system. The method for 

developing reasonable weight comparisons consists of building component models for finite-element 

analyses (FEA) of the exoskeletal engine system. The intention is to look at all major components in the 

engine system with some consideration given to mounting and assembly procedures for all rotating parts. 

Hopefully, this approach will unveil any potentially significant benefits as well as any possible technical 

hurdles pursuant to a new direction in turbine engine technology. Finally, the comparison of engine 

weights as a product of the investigation will serve as the principal metric for an evaluation of 

exoskeletal versus conventional engine benefits. 

Description of Exoskeletal System Concept 

 The exoskeletal engine architecture for this study is based on the Rolls-Royce (Allison) AE3007 

engine. The design dimensions were scaled from a sketch of a longitudinal section through the AE3007 

in Jane’s Aero-Engines (ref. 3). These scaled dimensions were input to NASA’s engine weight estimate 

computer code WATE (Weight Analysis Turbine Engine, refs. 4 and 5) to develop a government 

interpretation of an AE3007-like engine. Scaled dimensions and the output from the WATE code 

(appendix A) were used to generate a Pro/Engineer® (ref. 6) CAD model representation of the AE3007 

exoskeletal engine. 

 The AE3007 engine shown in figure 4 is a two-shaft subsonic turbofan, and the shafts are concentric. 

The AE3007 engine has an axial-flow, high-pressure compressor, a combustion chamber, and a highpressure turbine (HPT) all spinning on the outer shaft. The HPC has 14 stages, variable-inlet guide vanes, 

and variable-inlet stators in the first 5 rows. The overall pressure ratio of the HPC is 23. The combustion 

chamber is an annular design with 16 fuel nozzles. The HPT is a two-stage axial design with air-cooled 

blades. A three-stage, low-pressure turbine drives the inner shaft, which in turn drives a single-stage, 

wide-chord fan. The fan has a mass flow of 260 lb/sec and a bypass ratio of 5. The outer fan duct and jet 

pipe are composite designs. 
NASA/TM—2005-213369 10 

 Heritage of Rolls-Royce (Allison) AE3007-based NASA engine cycle model.⎯The NASA AE3007 

cycle model was developed to support NASA’s goals assessment (ref. 7). Analytical models of various 

airframes and propulsion systems representative of the current state of the art were developed in 

sufficient detail to allow the inclusion of new technologies to assess the benefit of these technologies on 

system performance (size, cost, emissions, maintainability, etc.). To disseminate study assumptions and 

results, NASA developed models of these airframes and propulsions systems, based on publicly available 

information and good engineering judgment. Although the models developed may be similar in overall 

performance and capability to the actual airframes and propulsion systems in use, there could be 

significant differences between the NASA-developed models and the actual systems. 

 Development of AE3007 thermodynamic model.⎯As part of that activity, a model of a 

50-passenger regional jet based on the EMB145 was developed, with its propulsion system based on 

the Rolls-Royce (Allison) AE3007. For its size and thrust class, the AE3007 is a moderately high-bypass, 

two-spool, mixed-flow turbofan. It uses only axial-flow components, reducing engine and core diameter. 

For this study, it was thought that its core diameter was within the state of the art for composite rotating 

materials. A one-dimensional, steady-state, thermodynamic cycle model was developed using the NASA 

Engine Performance Program (NEPP, ref. 8). A block model of the engine is shown in figure 5, 

indicating gas flow paths and mechanical connections. Table 1 shows gas conditions throughout the cycle 

at sea-level static conditions based on assumed component and overall engine performance (the actual 

NEPP output is listed in appendix B). Compressor and turbine maps were generated (refs. 9 and 10) 

based on component size and performance or were scaled from existing in-house maps with similar 

characteristics. The actual compressor and high-pressure turbine maps used are shown in figures 6 to 8. 

These maps are used for estimating the off-design performance of these components while the cycle 

model is “flown” over a range of Mach, altitude, and engine power conditions to generate performance 

data to be included in mission analyses. In addition, component maximum performance points are 

recorded (mass flow, temperature, pressure, torque, etc.) to be used to develop a flow-path weight model. 
NASA/TM—2005-213369 11 
NASA/TM—2005-213369 12 

TABLE 1.⎯NEPP AE3007 STATION PROPERTIES 

Station Mass flow, Pressure, Temperature, Fuel-to-air ratio 

number lbm/s psia °R 

1 257.4 14.7 545.7 0.0000 

2 257.4 14.5 545.7 

3 257.4 24.0 641.9 

4 41.7 24.0 641.9 

5 41.7 23.7 641.9 

 

6 36.1 353.4 1465.0 

7 36.1 348.1 1465.0 

8 36.8 332.5 2792.1 .0221 

9 41.5 72.4 1942.6 .0196 

10 41.5 72.0 1942.6 .0196 

 

12 257.3 21.0 775.7 .0031 

13 257.2 20.9 775.7 .0031 

14 257.2 20.9 775.7 .0031 

15 41.5 18.1 1426.0 .0196 

 

16 41.5 18.0 1426.0 .0196 

20 215.8 23.6 641.9 .0000 

21 4.8 353.4 1465.0 .0000 

22 215.8 23.6 641.9 .0000 
NASA/TM—2005-213369 13 
NASA/TM—2005-213369 14 

 Development of the AE3007 flow-path and weight model.⎯An engine flow-path and weight 

estimation model was developed using the WATE program (refs. 11 and 12) and was based on maximum 

conditions determined from NEPP. A graphical output detailing the WATE analysis of the engine flow 

path is shown in figure 9. The WATE code estimates the total weight of various components and 

subcomponents based on user inputs and conditions experienced. The weight of each component is 

developed from its subcomponent parts. For example, the compressor weight is determined on a stageby-stage basis, based on the flow conditions and performance of each particular stage. Input parameters 

include blade hub-to-tip ratio, aspect ratio, disk type, and compressor end support (e.g., frame and 

bearing type). Different materials may also be chosen for the various parts (disk, rotor and stator blades, 

surrounding case, etc.). Such detail was required for previous study efforts to assess the effect of 

technologies that might be applicable to only one part, such as blades, disks, and so forth. This level of 

detail was also helpful to start the analysis of an exoskeletal version of the AE3007. However, the WATE 

analysis does not go into detailed blade design, the development of which will be discussed the following 

section. Since the exoskeletal analysis was limited to the high-pressure spool of the engine, more detailed 

gas flow conditions (from NEPP/WATE) and mechanical data (from WATE) were used for subsequent 

analyses. Tables 2 and 3 show these data. 
NASA/TM—2005-213369 15 

TABLE 2.⎯NEPP/WATE AE3007 FLOW CONDITIONS THROUGH 

TURBOMACHINERY 

Component Stage Mass flow, Entrance Entrance Pressure 

 lbm/s temperature, pressure, ratio 

 oR psia across stage 

1 41.68 638 23.3 1.34 

2 699 31.2 1.30 

3 759 40.7 1.28 

4 820 52.0 1.26 

5 880 65.3 1.24 

6 940 80.7 1.22 

7 999 98.6 1.21 

8 1058 119.0 1.19 

9 1117 142.2 1.18 

10 1175 168.4 1.17 

11 1233 197.7 1.17 

12 1290 230.5 1.16 

13 1347 267.0 1.15 

14 1403 307.4 1.14 

HPC 

Exit 1465 351.9 ------ 

1 40.06 2693 329.2 2.01 

2 42.62 2319 163.5 2.27 

HPT 

Exit 42.62 1944 72.1 ------ 
TABLE 3.⎯WATE AE3007 STRUCTURAL WEIGHTS AND DIMENSIONS FOR 

HIGH-PRESSURE COMPRESSOR (HPC) AND HIGH-PRESSURE TURBINE (HPT)

Component Stage Length, Radius, r, in. Blade Weight, W, lbm Density, ρ, lbm/in.3 Pressure 

Lstage, Hub, Tip, Number, Height, Aspect Blade, Nuts, Disk, Rotor Acces- Disk, Blade, ratio, 

 in. rH rT NB HB, ratio, WB WN Wdisk drum, sories, ρdisk ρB PR 

 in. AR WRD WACC

HPCa 1 1.94 3.57 7.84 21 4.27 2.2 3.7 1.6 2.8 1.34  2 1.45 4.80 7.84 28 3.04 2.1 2.0 1.2 4.3 1.30  3 1.17 5.53 7.84 35 2.31 2.0 1.2 1.0 6.6 1.28  4 .97 6.02 7.84 42 1.82 1.9 .8 .8 10.9 1.26  5 .84 6.36 7.84 49 1.48 1.8 .6 .7 13.0 1.24  6 .73 6.62 7.84 56 1.22 1.7 .4 .6 13.0 1.22  7 .66 6.82 7.84 62 1.02 1.6 .3 .5 12.9 1.21  8 .60 6.97 7.84 68 .87 1.5 .2 .5 12.6 

0.284 

1.19 

 9 .56 7.09 7.84 74 .75 1.3 .2 .5 11.3 

0.160 

1.18 

 10 .53 7.19 7.84 78 .65 1.2 .3 .4 12.3 1.17  11 .50 7.28 7.84 82 .56 1.1 .2 .4 12.7 1.17  12 .49 7.34 7.84 84 .50 1.0 .2 .4 12.7 1.16  13 .48 7.40 7.84 85 .44 .9 .2 .4 12.9 1.15  14 .49 7.45 7.84 84 .39 .8 .2 .4 13.3 

.297 

.300 

1.14 

HPT 1 1.05 8.31 9.57 40 1.26 1.2 2.5 2.4 25.2 6.2 11.7 0.298 0.281 2.01  2 .67 8.31 9.51 63 1.20 1.8 1.5 1.5 20.5 3.9 7.4 .298 .281 2.27 a HPC blades have a constant taper ratio of 0.833. 

NASA/TM—2005-213369 1 6
NASA/TM—2005-213369 17 

Exoskeletal Engine High-Pressure Spool 

 Figure 10 shows the conceptual design of the exoskeletal drum rotor assembly. Inverting the AE3007 

design, the HPC and HPT rotor blades are mounted to the drum rotor and the stator rows are mounted to 

a stationary core. The stationary core outlines the rims of the original rotor disks on the AE3007 engine. 

Previous exoskeletal engine studies call for the complete elimination of the shafts and disks from the 

engine center (ref. 13), which would provide an open channel along the centerline (ref. 14). However, a 

stationary core with closed ends has been designed to support the stator blades. An AE3007-like 

exoskeletal design provides a convenient break at the LPT since it runs at a slower speed. Conceptually, 

the low-pressure spool would drive the fan via a separate drum rotor. The low-pressure spool drum rotor 

has been neglected in this study to focus on the exoskeletal high-pressure spool, which has greater 

physical demands. 

 A hypothetical exoskeletal engine assembly is depicted by the Pro/Engineer® model in figure 11, and 

the overall dimensions are shown in figure 12. To minimize the complexity of the study, three sections of 

the high-pressure spool (the HPC, combustion chamber, and HPT) were converted to an exoskeletal drum 

rotor design. The Pro/Engineer® model is intended for use in the conceptual design and analysis of an 

exoskeletal construction, and it is not a complete representation of the propulsion system. 

 Figure 11 shows the high-pressure spool as four sections. The HPC drum rotor assembly is broken 

down into two sections to permit material choice flexibility. Figure 13 shows the HPC drum rotor design 

in more detail. Graphite polyimide was investigated for the first seven stages where the temperature was 

below 600 °F. A titanium drum rotor is used for the last seven HPC stages where temperatures could rise 

to 950 °F. HastelloyTM is used in the combustion chamber and HPT drum rotors because gas temperatures 

could approach 2500 °F. 

 A bearing system is located at each end of the high-pressure spool to support the rotating drum. As 

shown in figure 11, the bearings transfer the loads between the drum rotor and the struts and permit 

relatively free rotation with minimum friction. 
NASA/TM—2005-213369 18 
NASA/TM—2005-213369 19 

 An examination of figure 13 shows that the exoskeletal engine assembly and maintenance will be 

more difficult than that of the conventional engine. One question is how the rotor blades will be 

assembled to the drum rotor using acceptable methods of construction. One answer was to divide the 

rotor into integral bladed rings for each stage, as shown in figure 14. Similarly, the stator would also 

consist of integral bladed-ring components. Since composite materials have an immense potential use in 

the rotor blades supported in compression, the integral bladed ring is a natural step to blend function and 

assembly. Manufacturing an integral bladed ring will also be a challenge, but this approach should reduce 

weight. To build the exoskeletal engine, each stage is installed in the rotor drum in an assembly 

procedure that alternates between stator bladed rings and rotor bladed rings. Each ring would provide the 

necessary space for the adjacent stage, or one could design a separate spacer. Although assembly and 

maintenance time would be increased, this concept would eliminate a longitudinal split of the drum rotor 

and longitudinal mounting flanges that would create rotor instability. 

 As stated earlier, the WATE computer code was used to generate dimensions and weights of 

exoskeletal components. HPC and HPT blade dimensions were also output from the WATE code. 

Table 4 shows the dimensions used to generate Pro/Engineer® models of each rotor stage. 

 A generalized NACA wing design was used to create the HPC rotor blade with stagger angles at the 

root and tip as shown in figure 15. The design of the HPT rotor blades is similar, but a camber angle of 

15° was used to define the HPT blade curvature as shown in figure 16. 
NASA/TM—2005-213369 20 

TABLE 4.⎯ BLADE DIMENSIONS FOR HPC AND HPT ROTOR STAGES 

Com- Radius, r, in. Hub Tip 

ponent 

Stage 

Inlet, 

rH

Exit, 

rT

Chord 

length, 

in. 

Blade 

width, 

in. 

Stagger 

angle, θ, 

deg 

Chord 

length, 

in. 

Blade 

width, 

in. 

Stagger 

angle, θ, 

deg 

Number 

of blades, 

NB

1 3.57 7.84 2.117 1.970 21.47 1.765 1.107 21 

2 4.8 7.84 1.587 1.374 30.02 1.322 .830 28 

3 5.53 7.84 1.273 1.041 35.09 1.061 .665 35 

4 6.02 7.84 1.056 .827 38.49 .880 .552 42 

5 6.36 7.84 .912 .690 40.86 .760 .477 49 

6 6.62 7.84 .802 .590 42.66 .668 .419 56 

7 6.82 7.84 .718 .516 44.05 .598 .375 62 

8 6.97 7.84 .655 .462 45.09 .546 .342 68 

9 7.09 7.84 .611 .425 45.93 .509 .319 74 

10 7.19 7.84 .577 .396 46.62 .480 .301 78 

11 7.28 7.84 .546 .370 47.25 .455 .285 82 

12 7.34 7.84 .535 .360 47.67 .446 .280 84 

13 7.4 7.84 .528 .352 48.08 .440 .276 85 

HPC 

14 7.45 7.84 .532 .353 48.43 .443 .278 

51.14 

84 

HPT 1 8.31 9.57 1.375 1.15 21.47 1.146 0.60 30 40 

2 8.31 9.51 .872 .73 26.54 .727 .38 30 63 
NASA/TM—2005-213369 21 

 

Structural Analysis and Assessment 

Blade Investigation 

Method of analysis.⎯Blade geometry was generated using the WATE code as summarized in the 

section Development of AE3007 flow-path and weight model. The blade height varied from 4.27 to 

0.39 in. between HPC stages 1 to 14, respectively. This variation in height for HPC blades is illustrated 

in figure 17. Parameters such as thickness, taper ratio, and chord length were varied during the 

investigation iteration process as required to achieve a positive margin of safety. In some cases, the 

camber angle was adjusted to increase compression stresses while decreasing tension stresses in the 

blade. A ceramic matrix composite (CMC) blade will perform better in compression than in tension. The 

NACA airfoil 65A010 was used as a representative blade section for the HPC and is shown in figure 18. 

 Using this approach, the AE3007-like blade geometries were derived and modeled using 

Pro/Engineer® to formulate the starting point for a finite-element analysis of each stage in the HPC and 

HPT. One blade was modeled for each stage in an exoskeletal design (blade in compression). The 

aerodynamic loads were assumed to be the same for the exoskeletal configuration as those for the 

conventional engine. The NEPP software code provided pressure loads for each stage, which was 

combined with rotational inertia for a combined static load case solved using MSC/NASTRANTM FEA 

software (ref. 15). The rotational speed of the engine is 16 400 rpm (overspeed condition). For each 

stage, the NEPP software code also provided temperature data, which was used to select appropriate 

materials for the finite-element analysis. Table 5 gives the entrance temperature range for stages 1 to 14 

of the HPC. 
NASA/TM—2005-213369 22 

 

TABLE 5.⎯ENTRANCE TEMPERATURES FOR 

HIGH-PRESSURE COMPRESSOR STAGES 

Stage Temperature, °F Stage Temperature, °F 

1 178 8 598 

2 239 9 657 

3 299 10 715 

4 360 11 773 

5 420 12 830 

6 480 13 887 

7 539 14 943 

High-pressure compressor analysis.⎯The first seven stages of the HPC utilize a lightweight 

polymer matrix composite (PMC) material. Graphite polyimide was selected as a suitable option having 

temperature-dependent properties up to 600 °F as shown in table 6. For stages 8 to 14, Titanium 6Al−4 

Valloy, hereinafter referred to as Titanium, was selected based on temperature capabilities up to 900 °F 

as shown in table 7. Other candidate materials used in the investigation of HPC blades include aluminum 

and WaspaloyTM with allowable stresses as shown in tables 8 and 9, respectively. These allowable 

material stresses were obtained from the Metallic Materials and Elements for Aerospace Vehicle 

Structures (ref. 16) and include the appropriate reductions for temperature. Similarly, the composite 

material properties (graphite polyimide) were derived from the HyperSizerTM (ref. 17) data base but with 

a 50-percent knockdown applied to account for life and durability effects. A summary of mechanical 

properties for materials in this investigation is presented in table 10. 

TABLE 6.⎯MATERIAL ALLOWABLES USED FOR GRAPHITE POLYIMIDE 

[Includes 50-percent knockdown for durability.] 

Temperature, Allowables 

°F Ultimate tensile force, Ultimate compressive force, Ultimate shear force, 

Ftu, Fcu, Fsu, 

 klbf/in.2 klbf/in.2 klbf/in.2

72 113 100 6.5 

600 105 51 3.1 
NASA/TM—2005-213369 23 

TABLE 7.⎯MATERIAL ALLOWABLES USED FOR TITANIUM 6AL−4V 

Temperature, Allowables 

°F Ultimate tensile force, Ultimate compressive force, Ultimate shear force, 

Ftu, Fcu, Fsu, 

 klbf/in.2 klbf/in.2 klbf/in.2

100 134 133 85 

200 126 119 80 

300 114 107 75 

400 103 96 70 

500 94 89 67 

600 92 88 65 

700 90 85 58 

800 86 82 52 

900 77 71 46 

TABLE 8.⎯MATERIAL ALLOWABLES USED FOR ALUMINUM 

Temperature, Allowables 

°F Ultimate tensile force, Ultimate compressive force, Ultimate shear force, 

Ftu, Fcu, Fsu, 

 klbf/in.2 klbf/in.2 klbf/in.2

100 66 58 40 

200 66 54 38 

300 51 44 30 

400 30 14 14 

500 25 8 7 

550 25 6 

600 25 5 

650 24 5 

700 24 

6 

4 

TABLE 9.⎯MATERIAL ALLOWABLES USED FOR WASPALOYTM

Temperature, Allowables 

°F Ultimate tensile force,a Ultimate compressive force,a Ultimate shear force,b

Ftu, Fcu, Fsu, 

 klbf/in.2 klbf/in.2 klbf/in.2

100 138 138 83 

200 136 136 82 

300 132 132 79 

400 129 129 77 

500 127 127 76 

600 125 125 75 

700 123 123 74 

800 121 121 73 

900 120 120 72 

1000 118 118 71 

1100 115 115 69 

1200 112 112 67 

1300 104 104 62 

1400 90 90 54 

1500 70 70 52 a

Applied 20-percent knockdown factor to allowable strength at 1⁄2-hr exposure. 

b

Assumed 60-percent ultimate strength as shear allowable. 
NASA/TM—2005-213369 24 

TABLE 10.⎯ SUMMARY OF MECHANICAL PROPERTIES FOR BLADE ANALYSIS 

Material Modulus of Material Coefficient of Poisson’s Allowables 

 elasticity, E , density, ρ, thermal expansion, ratio, Ultimate tensile Ultimate compressive 

 lb/in.2 lbm/in.3 α, μ force, Ftu, force, Fcu, 

 1/°F klbf/in.2 klbf/in.2

Graphite 2.4×107 0.056 2×10−5 0.33 See table 6 See table 6 

polyimide 

Titanium 1.48×107 .160 5.2×10−6 .3 See table 7 See table 7 

Aluminum 1×107 .097 1.3×10−6 .33 See table 8 See table 8 

WaspaloyTM 30.6×106 .298 7.6 to 7.9×10−6 .3 See table 9 See table 9 

 The solid model blade geometry for each stage of the HPC was imported in the MSC/PATRANTM

preprocessing software code (ref. 18) to create a finite-element model (FEM) for analysis. The initial 

blade models used a straight taper (ratio of tip chord to root chord is 1.0). Boundary conditions for 

conventional engine blades were fixed at the root or inside radius of the blade whereas the exoskeletal 

blade fixity is inverted as illustrated in figure 19. 

 Loads were applied to the exoskeletal blade FEM, including aerodynamic pressure acting on the 

blade combined with the rotating inertia. A static solution was then performed using MSC/NASTRANTM

for each stage of the HPC. The maximum radial displacement was 0.006 in. at the stage 1 blade tip. The 

maximum Von Mises stress in the first seven stages using graphite polyimide was 26.2 klb/in.2

 and 

occurred at the stage 5 blade root, resulting in a margin of safety of +0.95. The maximum buckling 

margin for the stage 1 PMC blades in compression is +2.24. It is interesting to note that both titanium and 

aluminum materials were investigated for the HPC stage 1 blades, and these selections fail in buckling 

for an exoskeletal design. Shear stress at the root of the stage 1 blade results in a negative margin of 0.60. 

Incorporating a taper in the blade design may reduce the negative margin, although a blade redesign such 

as this was not pursued in this investigation. 

 Stages 8 to 14 were investigated in a similar way, but the taper ratio of 1.0 was kept for each stage in 

this section of the HPC. Although Titanium 6Al−4V was the material of choice, a second material, 

WaspaloyTM, was also investigated for these stages. The applied loads were combined aerodynamic 

pressure and rotational inertia for the static solutions using MSC/NASTRANTM. The peak stress and 

deflection results for this section of the HPC blade investigation are given in table 11 along with buckling 

margins of safety. 
NASA/TM—2005-213369 25 

TABLE 11.⎯BLADE INVESTIGATION RESULTS FOR STAGES 1 To 14 

Stage Material Maximum stress, Radial displacement, Margin of safety, 

 σmax, in. M.S. 

 klb/in.2 (buckling) 

1 Graphite polyimide 25.3 0.006 +2.24 

5 Graphite polyimide 26.2 .0009 +9.18 

8 Ti6Al−4V 28.0 .0008 +4.22 

8 WaspaloyTM 34.4 .0005 +4.80 

9 Ti6Al−4V 28.3 .0006 +6.02 

 In summary, the lighter weight PMC works acceptably well in compression for the first seven stages 

of the HPC. Thus, graphite polyimide was used in tabulating the weights of the first seven stages of the 

compressor. Stages 8 to 14 of the HPC operate at temperatures above the practical range for PMC. 

Titanium was the selected material for blades in these stages and was used for weight calculation 

purposes. 

High-pressure turbine analysis.⎯The HPT two-stage turbine is part of the high-pressure spool 

located aft of the combustion chamber and is subjected to higher temperatures than the HPC. Similar to 

the HPC, the blade geometry for the HPT was conceptually designed by first scaling the AE3007 drawing 

with output from the conventional design and WATE code for specific dimensions, as shown in table 4. 

Thermal analysis was then used to establish a thermal distribution across the blade as shown in figure 20. 

The intent of the investigation was to select a material that would not require cooling of the HPT blades 

and thereby achieve greater weight efficiency in an exoskeletal design. Based on the service temperature 

regime, a silicon-carbide-fiber matrix composite (SiC/SiC) was the preferred material for strength-toweight properties. This CMC has the high-temperature capability and low-oxidation characteristics 

required for a combustion engine environment. 

 The assumed SiC/SiC properties used for this investigation are shown in table 12. The properties 

shown are estimated for a 1500 °F temperature under life-cycle service conditions. Although strength and 

durability in life-cycle applications are not readily available for SiC/SiC materials, the assumption was 

that these would be attainable values for current technology in CMC material systems. 
NASA/TM—2005-213369 26 

TABLE 12.⎯PROJECTED CMC PROPERTIES AND ALLOWABLES 

FOR EXOSKELETAL ENGINE ENVIRONMENT 

Modulus of elasticity, Allowables 

Mlb/in.2

E1,2 E3

Shear 

modulus,

G12, 

Mlb/in.2

Poisson’s 

ratio, 

μ12

Coefficient 

of thermal 

expansion, α, 

1/°F 

Compressive 

stress, Flc, 

klb/in.2

Tensile 

stress, Flt, 

klb/in.2

Shear, 

F12, 

klb/in.2

30.4 21.3 11.4 0.18 2.42×10−6 22.5 15.5 6.2 

 At the start of the HPT investigation, the first-stage blades had a height of 1.26 in. and a root chord 

length of 1.146 in. The applied loads consisted of aerodynamic pressure (165.7 psi) combined with a 

16 400-rpm inertial load. The blade model was rigidly supported at the root to simulate the integral blade 

ring held by the rotor drum for the inverted exoskeletal design. The second-stage HPT blades were 

1.20 in. in height and had a root chord length of 0.727 in. Similarly, the inertial loading plus pressure 

(91.5 psi) were applied as a static load on the blade. A camber angle for the HPT blades was set at 15° 

and the taper ratio at 0.80. The analysis results for the first attempt of these HPT blade configurations 

show high negative margins for compression, tensile, and shear stresses in the exoskeletal design. 

 Two steps were taken to alleviate the high negative margins for stress: 

 1. The root chord length was increased in half-inch increments. 

 2. The blade thickness was increased by 25 percent over the derived AE3007 values. 

These changes were effective in bringing the exoskeletal tensile and compressive stresses to a positive 

margin, but not until the chord length reached approximately 2.75 in. The practicality of this structuraldriven parameter on engine performance is not part of the investigation. The results of this parametric 

study of varying the chord length are shown for the exoskeletal design in table 13. As in the case of the 

compressor, shear margins were negative and will require further design improvement. It was observed 

that the material allowable reductions at the service temperature were affecting the margins of safety. 

Based on the findings of this investigation, the elimination of cooling in the HPT will not be as readily 

achieved as it was originally believed. If cooling can be done in an exoskeletal design, the allowable 

stresses would increase, which helps to achieve positive margins of safety. The addition of cooling in the 

blades is complex and may bring weight penalties into the design. 

TABLE 13.⎯EXOSKELETAL HIGH-PRESSURE TURBINE BLADE 

MARGINS OF SAFETY FOR VARIED CHORD LENGTH 

Chord 

length, 

in. 

Von Mises stress, 

σvm,

klb/in.2

Maximum principal 

stress (M.S. a), 

σ1, 

klb/in.2

Minimum principal 

stress (M.S. a), 

σ2, 

klb/in.2

Shear 

stress, 

στ , 

klb/in.2

1.146 167.0 62.7 (–0.75) –182 (–0.88) 90.8 (–0.93) 

1.719 37.5 33 (–0.53) –44.4 (–0.49) 20.2 (–0.69) 

2.0 29.1 29.2 (–0.47) –36.1 (–0.38) 15.4 (–0.60) 

2.5 19.9 20.8 (–0.25) –24.6 (–0.09) 10.5 (–0.41) 

3.0 15.9 12.4 (+0.25) –19.7 (+0.14) 8.2 (–0.24) 

a

M.S., margin of safety. 

 Using the temperature distribution shown in figure 20, a thermal case was added to the FEA in an 

attempt to understand the full environmental effects on turbine blades. Difficulties were encountered, 

however, in the region of the blade root where rigid fixity was assumed. This fixity produced very high 

stress at the blade root for the static thermal solution. This result is very conservative since rigid fixity at 

the interface of the blade and integral blade ring (not modeled) is not entirely true. There will be some 
NASA/TM—2005-213369 27 

distortion of the ring due to expansion that may relieve the high stress at the root. When the root 

constraints are removed, the thermal stress is no longer an issue. The real answer lies somewhere in 

between and would require a fully modeled blade-ring component to gain an understanding of the 

thermally induced stresses. This component would require a higher fidelity thermal analysis and perhaps 

an engine cycle analysis, which was beyond the scope of this investigation. Thus, thermal loads were not 

included in the analysis and margin calculations for this investigation. 

 Some benefits were observed for CMC material in an exoskeletal arrangement for the HPT blades. 

There is insufficient data, however, to conclude that negative margins in the exoskeletal design can be 

eliminated. Increasing the chord length lowers bending stresses at the root but may not be feasible for 

aerodynamic and flow-rate considerations. Similar to the HPC blades, shear margins are negative in the 

CMC turbine blades, which would require a design fix. Unresolved negative margins may be an issue in 

the use of SiC/SiC blade material for the HPT stages. The addition of cooling would increase the 

allowable stresses while adding complexity to the design, and this would not necessarily remove the 

negative margins for CMC material. Nonetheless, the CMC material was used for the tabulation of 

weight. It remains to be proven that CMC blades can be used in a high-pressure turbine stage meeting all 

strength and durability requirements. The purpose of this investigation, as stated before, was to take an 

aggressive approach for the potential exoskeletal engine design. 

Drum Rotor Investigation 

 This section summarizes the stress analysis and sizing of the drum rotor in the exoskeletal engine 

design concept. This work is part of an overall effort to determine a feasible size and weight for the 

exoskeletal drum rotor. The objective of this stress analysis is to determine the adequacy of the drum 

rotor concept when it rotates at 16 400 rpm, the engine overspeed condition. 

 This stress analysis is performed using the ANSYS® finite-element analysis software (ref. 19). A 

longitudinal section through the drum can be modeled as a two-dimensional axisymmetric model as 

shown in figure 21. The final stress state for this section as a body of revolution is then determined using 

ANSYS for analyzing an axisymmetric solid. In a rotating body, the two components of displacement 

(radial and axial) in any plane section of the body along its axis of symmetry completely define its state 

of stress. Thus, the use of an axisymmetric model greatly reduces the modeling and analysis time 

compared with that of an equivalent three-dimensional model. 

 The exoskeletal drum rotor is modeled on a 360° basis as shown in figure 13. Since ANSYS requires 

the model to be defined in the plane Z = 0.0, the global Cartesian Y-axis is assumed to be the axis of 

symmetry. Furthermore, the model is developed only in the +X-quadrants, which defines the radial 

direction. 

 To minimize the complexity of the study, this analysis focuses only on three sections of the highpressure spool of the exoskeletal rotor. The geometry is scaled from the AE3007 drawing, and the 

dimensions are shown in figure 22. The drum rotor is separated into four sections to permit material 

choice flexibility. The material choices are shown in figure 23. 
NASA/TM—2005-213369 28 

 Figure 24 shows the boundary conditions and loads applied to the exoskeletal drum rotor model. 

Since the rotor blades are discontinuous about the axis, an equivalent centrifugal force for each row of 

blades is applied to the inner radius of the disks. The weight of each rotor blade is obtained from the 

Pro/Engineer® model of the exoskeletal rotor. The total centrifugal force is calculated as a lumped mass 

at the disk inner radius spinning at 16 400 rpm: 

( ) 2

2 row 2

60 a a

W N

F mr r

g

⎡ π ⎤

= ω= ⎢ ⎥ ⎣ ⎦

 (1) 

where F is the total centrifugal force, m is lumped mass, ra is the disk inner radius, ω is the rotational 

speed (radians per second), Wrow is the total weight of all of the blades in that row, g is the gravitational 

constant, and N is the rotational speed (revolutions per minute). These centrifugal forces are applied to 

the FEM on a 360° basis, which evenly distributes the blade load on each disk. 

 Boundary conditions are applied to the drum rotor model at the bearing locations. There are two 

radial bearings (one at each end), and one thrust bearing at the inlet to the high-pressure spool. Referring 

to figure 24, an axial constraint is applied at the thrust bearing location. This axial constraint prevents 

free body motion along the spin axis. 

 The radial bearings are neither infinitely compliant nor rigid. Without accurate stiffness values, the 

actual bearing stiffness falls between no constraint and a rigid constraint. To bound the problem, two 

model cases are defined in the initial investigation: (1) rigid constraints are applied at the two radial 

bearing locations (as shown in fig. 24) and (2) no radial constraints are applied at the ends of the drum 

rotor model (as shown in fig. 25). 
NASA/TM—2005-213369 29 

Case 1.⎯Figure 25 shows a radial displacement contour plot for the exoskeletal drum rotor with 

radial constraints. The displacement plot shows zero radial growth at the rotor ends and a maximum 

radial displacement of 0.092 in. at the peak of the combustion section shell. The radial stress contours for 

case 1 are shown in figure 26. The magnitude of the peak radial stresses indicates that the model is likely 

to be overconstrained at the aft radial bearing location. 

 The tangential stress contours for case 1 are shown in figure 27. The maximum tangential stress 

value of 314 720 psi at the end of the combustor section may be too great because the drum rotor is 

overconstrained. If the drum rotor is not free to grow in the radial direction at the ends, the heavy 

combustor section increases the stress near the joint flange at the high-pressure turbine. 

 The minimum and maximum axial stresses shown in figure 28 also indicate that the model is 

overconstrained at the aft radial bearing location. 

Case 2.⎯Figure 29 shows the radial displacement contour plot for the drum rotor without radial 

constraints. The displacement plot shows radial growth at the rotor ends and a maximum radial 

displacement of 0.092 in. at the peak of the combustion section shell. The maximum radial displacement 

in case 2 is similar to case 1. However, each end of the drum rotor assembly is free to grow in the radial 

direction. 
NASA/TM—2005-213369 30 
NASA/TM—2005-213369 31 

 The radial stresses for case 2 are shown in figure 30. The radial stresses in case 2 are lower because 

the drum rotor is free to grow in the radial direction at the ends. 

 The tangential stress contours for case 2 are shown in figure 31. Now, the maximum tangential stress 

has dropped to a value of 263 610 psi. The maximum tangential stress in the combustor section near the 

joint flange is lower because the drum rotor is free to grow in the radial direction at the ends. 

 The minimum and maximum axial stresses, as shown in figure 32, also occur in the combustor 

section near the joint flange with the HPT. These axial stresses appear to be bending stresses because the 

heavy combustor section hinges at that joint flange. 
NASA/TM—2005-213369 32 

 As a check to the validity of peak stresses observed in the drum rotor, a simplified hand calculation 

of a rotating disk as shown in figure 33 was used to determine the stresses due to rotation. Additional 

forces operate on the disks because the rotor blades are also pushing out on the inner radius of the disks, 

which cause additional stresses at that interface. There are no additional forces on the combustion 

section, which makes it easier to compare the finite-element results with fundamental theory. 
NASA/TM—2005-213369 33 

 In a spinning disk, the inertia forces create tangential and radial stresses due to the tangential 

elongation of the disk. In other words, both the circumference and the radius of the disk grow due to the 

centrifugal forces acting on material within the spinning disk. 

 From Shigley (ref. 20), the maximum radial stress for the simplified rotating disk occurs at the radius 

location r = (rarb)

1⁄2: 

( )

( ) ( ) ( )

2 2

,max

2

2

3

8

3 0.33 16 400 2 0.298 10.375 9.75

8 386.4 60

370 psi

r ba r r

g

+ μ γω σ= −

+ π ⎡ ⎤ = − ⎢ ⎥ ⎣ ⎦

=

 (2) 
NASA/TM—2005-213369 34 

where σr,max is maximum radial stress, μ is Poisson’s ratio, γ is weight density, ω is rotational speed, g is 

the gravitational constant, and rb and ra are the outer and inner disk radii, respectively, 

 Referring to figures 26 and 30, this stress value certainly falls within the stress contours shown in 

each plot. This radial stress calculation cannot validate either case 1 or case 2. 

 From Shigley (ref. 20), the maximum tangential stress σt,max occurs at the inner boundary of the disk 

where r = ra. For the combustion section this value is calculated as follows: 

( ) ( )

2

2 2

,max

2

2 2

3 1

4 3

3 0.33 16 400 2 0.298 1 0.33 10.375 9.75

4 386.4 60 3 0.33

240 057 psi

t ba r r

g

+ μ γω − μ ⎛ ⎞

σ= + ⎜ ⎟ ⎝ ⎠ + μ

+ π ⎡ ⎤ ⎛ ⎞ − = + ⎢ ⎥ ⎜ ⎟ ⎝ ⎠ + ⎣ ⎦

=

 (3) 

This stress calculation is in close agreement with the finite-element results for the exoskeletal drum rotor 

without radial constraints at the bearings. 

 Encouraged by these calculations, another study is performed using the model from case 2 (no rigid 

constraints at the radial bearing locations). Although the stresses were significantly reduced, the 

maximum tangential stress in the exoskeletal drum rotor exceeds the ultimate strength of the material, 

Hastelloy XTM, selected for the shell over the combustion section. An adequate design is needed before 

the weight of the exoskeletal rotor concept is tallied. 

Trade study.⎯The normal procedure for performing a finite-element analysis involves (1) defining 

the model, its boundary conditions, and loads, (2) obtaining a solution, and (3) interpreting the results. If 

the results indicate a design change is necessary, the whole process must be repeated. The model 

definition alone may be time consuming. The ANSYS Parametric Design Language (APDL, ref. 21) 

gives the user the ability to automate this process with user-defined variables (parameters) and 

programming input. The exoskeletal drum rotor model was generated using design parameters that could 

be easily changed. The ANSYS input data file for the exoskeletal drum rotor is found in appendix C. 

 The exoskeletal drum rotor model is generated using design parameters to vary the thickness of the 

drum rotors and rotor disks (i.e., bladed rings). A trade study is performed with this model to track the 

impact on rotor stresses as these thicknesses are varied. The objective of this study is to find a design in 

which the stresses are below the material limits while minimizing the rotor weight. A range of plausible 

drum rotor and bladed-ring thicknesses are used in this study to minimize stresses and weight. 

 Figure 34 shows the result of this trade study in a carpet plot. Unfortunately, this study shows that no 

combination of drum rotor and bladed-ring (disks) thicknesses can overcome the inertia forces in this 

design concept. Referring to the stress equations for a rotating disk, the maximum tangential stress 

(eq. (3)) is proportional to the square of the radius. The stress in the exoskeletal rotor increases rapidly as 

the rotor diameter and thicknesses increase. By moving the rotating body out to a larger diameter, the 

exoskeletal rotor concept increases the mass moment of inertia and weight beyond the capability of 

current materials. The maximum tangential stress is also proportional to the square of the rotational 

speed. Reducing the engine speed would significantly reduce stresses in the rotor. Furthermore, since the 

maximum tangential stress is directly proportional to weight density, the use of high-strength-to-weight 

composite materials would help to reduce stresses. However, these composite materials may not be 

suitable for high-temperature combustion environments where the strength is needed. 

 The weight estimates in this study will use a bladed-ring (disk) thickness of 0.375 in. and a drum 

rotor thickness of 0.125 in. The resulting weights generated in tables 14 to 16 are considered a lower 

bound for the structural weight of a drum rotor assembly in this size class. 
NASA/TM—2005-213369 35 

TABLE 14.⎯HIGH-PRESSURE COMPRESSOR DRUM ROTOR WEIGHT ESTIMATE 

(a) Blade ring assembly 

Rotor Ring Single Number Total blade Total blade-ring 

stage Mass, Weight, airfoil of blades airfoil weight, assembly weight, 

 lb-sec2/in. lb weight, lb lb 

 lb 

1 9.841×10−3 3.8025 5.83×10−2 21 1.223 5.03 

2 6.031×10−3 2.3302 2.38×10−2 28 .667 3.00 

3 6.469×10−3 2.4997 1.18×10−2 35 .412 2.91 

4 5.592×10−3 2.1608 6.41×10−3 42 .269 2.43 

5 4.962×10−3 1.9172 3.90×10−3 49 .191 2.11 

6 4.879×10−3 1.8854 2.49×10−3 56 .139 2.02 

7 3.975×10−3 1.5358 1.67×10−3 62 .103 1.64 

8 9.242×10−3 3.5710 3.38×10−3 68 .230 3.80 

9 1.089×10−2 4.2065 2.54×10−3 74 .188 4.39 

10 1.010×10−2 3.9039 1.96×10−3 78 .153 4.06 

11 9.633×10−3 3.7223 1.51×10−3 82 .124 3.85 

12 1.049×10−2 4.0552 1.30×10−3 84 .109 4.16 

13 8.459×10−3 3.2684 1.11×10−3 85 .095 3.36 

14 9.242×10−3 3.5710 1.00×10−3 84 .084 3.66 

(b) Drum rotor 

Drum rotor Flange 

Stage Mass, 

lb-sec2

/in. 

Weight, 

lb 

Number Mass, 

lb-sec2/in. 

Weight, 

lb 

Total drum 

rotor weight, 

lb 

1 2.247×10−2 8.6840 1 1.298×10−2 5.0168 13.70 

2 4.441×10−2 17.1589 -- -------------- --------- 17.16 

Total drum rotor weight, lb 77.28 
NASA/TM—2005-213369 36 

TABLE 15.⎯COMBUSTOR DRUM ROTOR WEIGHT ESTIMATE 

Combustor drum rotor Flange 

Stage Mass, 

lb-sec2

/in. 

Weight, 

lb 

Number Mass, 

lb-sec2/in. 

Weight, 

lb 

Total drum 

rotor weight, 

lb 

3 1.112×10−1 42.9719 --- -------------- --------- 42.97 

Total combustor drum rotor weight, lb 42.97 

TABLE 16.⎯HIGH-PRESSURE TURBINE DRUM ROTOR WEIGHT ESTIMATE 

(a) Blade ring assembly 

Rotor Ring Single Total blade Total blade-ring 

stage Mass, Weight, airfoil airfoil weight, assembly weight, 

 lb-sec2/in. lb weight, lb lb 

 lb 

Number 

of blades 

NB

 

1 4.332×10−3 1.6737 ------ ----- 2.5 4.17 

2 1.123×10−2 4.3394 ------ ----- 1.5 5.84 

 

(b) Drum rotor 

Drum rotor Flange 

Stage Mass, 

lb-sec2

/in. 

Weight, 

lb 

Number Mass, 

lb-sec2/in. 

Weight, 

lb 

Drum rotor 

weight, 

lb 

4 4.498×10−2 17.3812 ----- -------------- --------- 17.38 

Total drum rotor weight, lb 27.39 

Rotor Dynamics Assessment 

 This section covers a preliminary lateral vibration analysis of a flexible beam designed to simulate 

the exoskeletal rotor design. The purpose of the analysis is to determine if this structure has natural 

vibration modes within the operating speed range of the AE3007 turbine engine. 

 In this analysis, the drum assembly is mounted to two flexible bearings, which in turn are rigidly 

fixed. Dynamic coupling with other systems in the engine through the bearings is neglected. Since 

bearings play a vital role in the dynamic behavior of a rotor, one must select the appropriate bearing 

stiffness to ensure against potentially destructive vibrations. Thus, another purpose of this analysis is to 

evaluate the bearing stiffness for the exoskeletal drum rotor design. 

 A computer code, CRTSPD (ref. 22), has been developed especially for determining the critical 

speeds and mode shapes of shaft systems. This program uses a lumped-parameter rotor representation to 

determine all critical speeds of the exoskeletal drum rotor design. A critical speed is defined as the 

condition where the rotating speed of the rotor coincides with the natural frequency (ref. 23). 

 Prudent design practice should prevent any coincidence of critical speeds and the rated speed. As a 

result, the following criterion has been adopted. For a single-span rotor, the critical speeds must be 

outside a range of 15 percent below to 25 percent above the running speed to provide an adequate margin 

of safety. The high-pressure compressor and turbine in the AE3007 engine have an operating speed of 

14 600 rpm, and a maximum overspeed condition of 16 400 rpm. Therefore, based on these running 

speeds, any critical speed must be below 12 410 rpm and above 20 500 rpm. 
NASA/TM—2005-213369 37 

Lumped-parameter vibration model.⎯Rotating motion requires a distinctive rotor structuraldynamic model and an eigenvalue analysis. Eigenvalue analysis refers to the mathematical calculation of 

undamped natural frequencies (critical speeds) and mode shapes. The distribution of the mass and 

stiffness along the rotor and the bearing support stiffness determine the mode shapes (ref. 23). 

 Structural-dynamic models for flexible rotors are derived by using Timoshenko’s beam models for 

bending and shear deflection (ref. 24), transverse and rotary inertia, and gyroscopic effects. The lumpedparameter vibration model is traditionally used to account for the distributed elastic and inertial 

properties of a rotor. The rotor is modeled as a collection of several rigid bodies connected by massless 

elastic beam elements. Assuming the rotor is axisymmetric, the general equation of motion for a lumpedparameter model, shown in figure 35, is defined by the following set of transfer matrix equations (ref. 2): 

( )

( )

2

2

1 0 00

0 1 00

0 10

0 01

n n

n n

n n n n

n n n n

V V m K

M M I J

⎡ ⎤ ⎧⎫ ⎧⎫ δ δ ′ ⎢ ⎥ ⎪⎪ ⎪⎪ θ θ ′ ⎨⎬ ⎨⎬ = ′ ω − ⎪⎪ ⎪⎪

′ ⎩⎭ ⎩⎭ − ω ⎣ ⎦

 (4a) 

3 2

1

2 1

1

1

1 62

01 2

00 1 0

00 1

n nn nn n n

n n n n nn

n n

n n n

L L EI L EI

L EI L EI

V V

M M L

+

+

+

+

⎧ ⎫ ⎧⎫ δ δ ⎡ ⎤ ′ ⎢ ⎥ ⎪ ⎪ ⎪⎪ θ θ′ ⎨ ⎬ ⎨⎬ = ′ ⎪ ⎪ ⎪⎪

′ ⎩ ⎭ ⎩⎭ ⎣ ⎦

 (4b) 

Where δ is radial deflection (inches), θ is angular deflection (radians), V is shear force (pounds), M is 

bending moment (pounds-inches), K is matrix stiffness (pounds per inches), I is area moment of inertia 

(inches4

), J is polar moment of inertia (inches4), L is shaft section length (inches), E is modulus of 

elasticity (pounds per square inch), and ω is shaft speed (radians per second). 

 Historically, two approaches have been developed to derive and solve the governing matrix equations 

for a lumped-parameter flexible rotor model (ref. 25). M.A. Prohl (ref. 26) and N.O. Myklestad (ref. 27) 

independently developed the transfer-matrix formulation method, which is now known as the MyklestadProhl method. Myklestad and Prohl both developed a tabular method to find the modes and frequencies 

of structures, such as airplane wings and turbine blades. C. Biezeno and R. Grammel developed (1959) 

the general mass-stiffness-matrix approach (ref. 25). Both these methods have been used to develop 

computer programs that define flexible-rotor matrix models from rigid-body equations and to perform the 

eigenvalue analysis. 
NASA/TM—2005-213369 38 

 Calculations of rotor critical speeds and the corresponding amplitudes of vibration require the use of 

a computer program such as CRTSPD. To analyze the complex geometry of the exoskeletal design, the 

rotor was divided into a suitable number of sections defining a series of interconnected uniform beams. 

 Figure 36 shows a schematic of the CRTSPD model, which can contain up to 30 sections. The 

computer input data for each section consist of an external weight, section length, shaft outer and inner 

diameters, transverse moment of inertia, and rotational (polar) moment of inertia. 

 The generic representation of the exoskeletal rotor is a single-span, two-bearing system. In this 

system, the rotor, bladed disks, and bearings are assumed to be axisymmetric. In addition, all materials 

are assumed to be quasi-isotropic. 

 The computer program CRTSPD allows the user to (1) model the disk as an integral part of the shaft 

or (2) model only the shaft and add the disk mass properties externally. The advantage to modeling the 

disks as integral members of the shaft is that the program automatically calculates the cross-sectional 

properties. The disadvantage is that the section stiffness may be more than twice that of the shaft under 

the disk. However, when the disk mass properties are put separately, one loses the additional stiffness 

that the disks add to the shaft. In the exoskeletal case, only the drum rotor is modeled with the disk mass 

properties applied externally. Since critical speeds (or natural frequencies) in a flexible system are 

proportional to stiffness (i.e., ωn ≈ (k/m)

1/2, this approach is more conservative. The exoskeletal rotor 

should be considerably stiffer than the original AE3007 because of its larger diameter. For this 

investigation, it is assumed that the exoskeletal rotor will behave as a rigid body on two soft bearings. If 

any critical speeds occur in the range of 12 410 to 20 500 rpm, the model must be modified to include the 

disk. 
NASA/TM—2005-213369 39 

 Table 17 shows the CRTSPD input data file for the exoskeletal rotor, and table 18 shows the bearing 

stiffness for each bearing used in this analysis. A bearing stiffness of 20 000 lb/in. is selected for this 

rotor (Andrew J. Provenza, March 19, 2003, NASA Glenn Research Center, Cleveland, OH, personal 

communication), which is a typical stiffness value for magnetic bearings. 

TABLE 17.⎯CRTSPD LUMPED-PARAMETER MODEL FOR EXOSKELETAL ROTOR 

CRTSPDM – MODAL ANALYSIS VERSION 2/2/77 

N = 30 STA. NB = 2 BRG. NSK = STIFFNESSES G = 0.0 NGYR = 1 MASS = 0 NPLOT = 2 

STATION WEIGHT SHAFT DIAMETER 

LB 

LENGTH IN. 

OUTSIDE INSIDE 

POLAR MOM. 

LB-IN.**2 

TRANS. MOM. 

LB-IN.**2 

Ex10-6 

LB/IN.**2 

DENSITY 

LB/IN.**3 

1 34.792 2.920 16.680 16.430 0.133E+05 5637. 24.00 0.056 

2 5.987 2.340 16.680 16.430 0.129E+04 643.1 24.00 0.056

3 3.870 2.440 16.680 16.430 0.128E+04 640.0 24.00 0.056

4 3.726 2.040 16.680 16.430 0.128E+04 638.0 24.00 0.056

5 3.153 1.930 16.680 16.430 0.127E+04 634.8 24.00 0.056

6 2.793 1.820 16.680 16.430 0.127E+04 633.3 24.00 0.056

7 3.090 1.420 16.680 16.430 0.129E+04 646.6 16.00 0.160 

8 3.060 1.310 16.680 16.430 0.132E+04 658.6 16.00 0.160

9 9.514 1.370 16.680 16.430 0.132E+04 657.7 16.00 0.160

10 5.758 1.260 16.680 16.430 0.131E+04 656.8 16.00 0.160

11 5.365 1.250 16.680 16.430 0.131E+04 654.7 16.00 0.160

12 5.171 1.290 16.680 16.430 0.131E+04 655.2 16.00 0.160

13 5.445 1.180 16.680 16.430 0.131E+04 653.9 16.00 0.160

14 4.587 1.180 16.680 16.430 0.130E+04 652.0 16.00 0.160

15 4.757 0.930 16.680 16.430 0.129E+04 647.5 16.00 0.160

16 5.454 0.000 16.680 16.430 0.125E+04 626.4 16.00 0.160

17 3.144 1.000 17.000 16.750 70.1 35.11 29.80 0.297 

18 3.063 2.000 17.950 17.700 235. 118.4 29.80 0.297

19 4.306 2.000 19.220 19.970 368. 185.5 29.80 0.297

20 4.261 1.720 20.400 20.150 412. 207.3 29.80 0.297

21 3.266 1.000 21.260 21.010 347. 173.9 29.80 0.297

22 1.233 0.000 21.580 21.340 138. 68.92 29.80 0.297

23 2.477 2.000 21.360 21.110 279. 140.4 29.80 0.297

24 4.903 2.000 20.930 20.680 542. 272.5 29.80 0.297

25 5.553 2.640 20.430 20.180 585. 295.0 29.80 0.297

26 7.936 0.000 20.140 19.900 322. 162.9 29.80 0.297

27 5.912 0.950 20.020 19.770 0.226E+04 1129. 29.80 0.297

28 6.954 1.450 20.020 19.770 0.243E+04 1213. 29.80 0.297

29 9.042 1.310 20.020 19.770 0.247E+04 1234. 29.80 0.297

30 28.350 0.000 20.020 19.770 0.720E+04 3598. 29.80 0.297

 ------- ------- --------- 

 196.920 42.750 48308.500 
NASA/TM—2005-213369 40 

TABLE 18.⎯CRTSPD BEARING STIFFNESSES 

BEARING STATIONS 1 3O 

SPEED DEPENDENT BEARING CHARACTERISTICS 

BEARING 

NUMBER 

BEARING 

LOCATION 

K1 

LB/IN 

K2 

LB-SEC/IN 

K3 

LB-SEC2/IN 

C1 

LB/IN 

C2 

LB-SEC/IN 

C3 

LB-SEC/IN 

STIFFNESS CASE NO. 1 

1 1 20000.0 0.0000E+00 0.0000E+00 0.0000E+00 28.50 0.0000E+00 

2 30 20000.0 0.0000E+00 0.0000E+00 0.0000E+00 28.50 0.0000E+00 

Results of eigenvalue analysis.⎯The modes of vibration (critical speeds) of the rotor system may be 

divided into two categories: (1) modes where the rotor behaves essentially as a rigid body and (2) 

flexural modes where the rotor whirls with a characteristic deflection curve. 

 Figure 37 shows the two rigid-body modes for the exoskeletal rotor. In the first critical speed at 

2467 rpm, the bearing support points at each end of the rotor move in phase with one another. The locus 

of the rotor’s whirling motion traces a cylindrical path. 

 At the second rigid-body mode, at a critical speed of 5776 rpm, the bearing support points move out 

of phase with one another. As seen in figure 37, there is a node point near or at the midspan where the 

deflection crosses the spin axis. The locus of the rotor’s whirling motion traces two cones, point to point. 

 Figure 38 shows the third mode of the exoskeletal rotor, which occurs at 261 570 rpm. The third 

critical speed is the first flexural mode of the exoskeletal rotor, and it is due to the bearing flexibility. 

Using differential equations, F.C. Linn and M.A. Prohl (ref. 28) showed that the rigid-body critical 

speeds both approach zero as the bearing flexibility approaches infinity, as shown in figure 39. In this 

case, the third critical speed at 261 570 rpm is the first “free-free” mode of the rotor system because there 

is no apparent restraint at either support point (ref. 29). 

 
NASA/TM—2005-213369 41 

Rotor dynamics summary.⎯Assuming the bearing stiffness for this rotor design is 20 000 lb/in. for 

the entire speed range, the first and second undamped critical speeds or rigid-body modes occur at 2467 

and 5776 rpm, respectively. These speeds are well below the operating speed of 14 600 rpm for the 

exoskeletal rotor design concept. Usually, the rotor flexural modes require more design consideration 

because the rotor and stator can vibrate out of phase with one another, which may cause rubbing of seals 

or internal damage. However, the third critical speed (or first flexural mode) occurs at 261 570 rpm, and 

this speed is well above the 16 400 rpm overspeed condition. Further study is recommended to analyze 

the response to imbalance, bearing location, bearing damping, and damping in the supporting structure. 

In this preliminary analysis, the limits placed on the critical speeds ensure satisfactory operation at the 

design speeds of the exoskeletal engine. 

Stationary Core Investigation 

Description.⎯The stationary core at the center of the exoskeletal engine forms the inner flow-path 

surface of the inlet, compressor, combustion chamber, turbine, and nozzle exit, whereas the drum rotor 

forms the outer flow-path surface. Additionally, the core functions as a backup structure for the stator 

guide vanes. The guide vanes are nonrotating, fixed-position blades integrally machined as a bladed-ring 

assembly. There is one guide vane ring assembly for each stage of the compressor and turbine. The ring 

assembly slides onto the stationary core with interlocking spacers between each stage row to 

accommodate the respective rotating stages of the drum. 

 The stationary core consists of four distinct sections along its length with each section made from a 

different material. The use of multiple materials gave the lightest weight design possible while providing 

adequate structural strength and thermal management. The first section enveloped the inlet and first 

seven stages of the compressor and was made of graphite polyimide to handle operating temperatures up 

to 600 °F. Titanium was used for the second section, which included the final seven stages of the 

compressor and just up to the combustion chamber. Combined pressure and temperature were the design 

drivers for the material selection in the combustion chamber. For this section, Hastelloy XTM was chosen 

for its high-strength and extreme-temperature properties. A CMC of SiC/SiC was selected for the last 

section, which consisted of the turbine and exit region assumed to be operating at 2200 °F or higher. This 

CMC does not have a particularly favorable strength-to-weight ratio but was selected to eliminate the 
NASA/TM—2005-213369 42 

need for cooling and thereby realize a net weight savings. Specific details of attaching the various shells 

to one another were not addressed in this study, but additional weight to account for attachments was 

added to the weight summary prediction. 

Analysis.⎯A finite-element shell model of the core structure was developed with appropriate 

material properties assigned at each section of the core. The model was constrained axially at the forward 

thrust ring station near the inlet. A second constraint was added at the aft ring location of the core to 

allow for lateral support of the structure and thermal growth along the axis. The applied boundary 

conditions provided an effective relief of any thermal expansion in the core structure. The primary load 

comes from the combustion pressure and resulting thrust of the engine. The load path for this pressure is 

in the compression of the core shell, and the thrust is transferred through attachment vanes between the 

stationary core and the nacelle at the thrust ring location. The net thrust causes an overall compression 

load in the core structure that could result in buckling of the shell. 

 Strength and linear buckling loads were determined for the stationary core. Adjustments to shell 

thickness were made to ensure positive margins of safety for strength and stability. A hand calculation 

was performed using the line loads to compute shell buckling. Results of this preliminary assessment 

yielded a thickness requirement of 0.05 in. for the graphite polyimide, 0.09 in. for the titanium, 0.12 in. 

for the HastelloyTM, and 0.12 in. for the SiC/SiC. These thicknesses were used in determining weight 

estimates for the stationary core assembly and guide vanes; the subtotal is 100.5 lb as listed in table 19. 

Figure 40 shows the stationary core Von Mises stress results for the core segments. Figure 41 illustrates 

the deformed shape of the first buckling mode. As shown, the critical section of the core has a 17-percent 

margin of safety against buckling for the shell thickness provided. 

TABLE 19.⎯EXOSKELETAL CONCEPT STATIONARY CORE 

[Ring thickness, 0.10 in.] 

(a) Weight summary 

Rotor Single Number Total blade Radius, Weight Axial location Ring 

stage airfoil of airfoil r, density, x0, x1, weight, 

 weight, blades, weight, in. γ, in. in. lb 

 lb NB lb lbm/in.3 

HPC 

1 0.0583 21 1.22 3.57 0.056 2.25 3.59 0.17 

2 .0238 28 .67 4.80 .056 4.72 5.79 .18 

3 .0118 35 .41 5.53 .056 7.24 8.15 .18 

4 .0064 42 .27 6.02 .056 9.28 10.19 .19 

5 .0039 49 .19 6.36 .056 11.31 12.01 .16 

6 .0025 56 .14 6.62 .056 13.19 13.78 .14 

7 .0017 62 .10 6.82 .056 14.59 15.23 .15 

8 .0034 68 .23 6.97 .160 16.02 16.41 .27 

9 .0025 74 .19 7.09 .160 17.37 17.80 .31 

10 .0020 78 .15 7.19 .160 18.60 19.09 .35 

11 .0015 82 .12 7.28 .160 19.86 20.32 .34 

12 .0013 84 .11 7.34 .160 21.10 21.66 .41 

13 .0011 85 .09 7.40 .160 22.38 22.74 .27 

14 .0010 84 .08 7.45 .160 23.55 23.92 .28 

HPT 

1 0.0080 40 0.32 8.31 0.076 39.04 39.95 0.36 

2 .0037 63 .23 8.31 .076 41.40 42.25 .34 

(b) Weight totals 

Blade airfoil, 

lb 

Ring, 

lb 

Blade and rings, 

lb 

Stator shell, 

lb 

Attachments, 

lb 

Total weight of blade 

and rings, stator shell, 

and attachments, 

lb 

4.54 4.09 8.6 88.9 3.0 100.5 
NASA/TM—2005-213369 43 

 

Bearing Subsystem 

 The exoskeletal engine concept requires large-diameter radial bearings on the outside diameter of the 

rotating drum located at both ends of the spool (as shown in fig. 42). A thrust bearing is located near the 

outside diameter at the inlet side of the compressor spool. The large diameter coupled with the high 

rotational speed poses significant challenges for bearing technology. State-of-the-art bearings for 

conventional engines with lubricating systems can operate at very high revolutions per minute and not 

exceed state-of-the-art rotational speed, which is approximately 4 M DN (ref. 30). For the case of the 

exoskeletal high-pressure spool, the bearings operate at 7 M DN, well beyond the state of the art. Recent 

work of Sullivan at the NASA Glenn Research Center has shown thermal problems with carbon-carbon 

(C-C) bearings for the exoskeletal application (ref. 31). There are no known lubricated systems that can 

handle this magnitude of velocity; consequently, noncontact bearing systems are needed to accommodate 

the demands of the exoskeletal high-pressure spool heat loads. Among the candidates for investigation 

would be (1) foil bearings and (2) magnetic bearings. 

Foil bearings.⎯The foil bearings shown in figure 43 are noncontacting and ride on a thin film of air, 

which is generated hydrodynamically (ref. 32) by the rotational speed. Lightweight foils are used to 

suspend and center the shaft. The current state-of-the art size for this type bearing is about 4 in. in 

diameter (ref. 31). Typical foil bearing applications use a single foil rolled around the shaft as seen in the 

figure. 

 Competing foil bearing design technology employs a bumped foil to suspend the shaft (fig. 44). For 

the large-diameter application on the exoskeletal engine architecture, a hybrid system is envisioned as 

one in which the bumped foil provides stiffness and multiple foils are used to keep the large-diameter 

shaft centered (as shown in fig. 45). The complexity of the design and the significant increase in diameter 
NASA/TM—2005-213369 44 

both lead to a significant and long-duration technical effort. Some of the drawbacks for the foil system 

include the high startup torque, the need for set-down/lift-off mechanical bearings and associated 

positioning hardware to accommodate anticipated duty cycle requirements. Unfortunately, no current 

bearing system can handle the heat generated by this system; consequently, an alternative system must be 

considered. For the purpose of this study, high-pressure air generated with an auxiliary power unit (APU) 

was selected to provide stiffness. With the help of a fast-acting control system, the bearing could be 

stiffened to handle extreme loads experienced in flight. The use of localized stiffening creates a viable 

backup bearing option, without relying on high-temperature material technology that may not be 

available in the near future. The weight estimation for this type of bearing system in the exoskeletal 

concept for this investigation was extrapolated since no data exist at the large diameter required. A 

concern about bearing stability is addressed by extending the width of the bearing to 4 in. 

 

 
NASA/TM—2005-213369 45 

Magnetic bearings.⎯In the case of a magnetic bearing system, there are some advances in this area 

of research and development (R&D). Most of this R&D, however, has been focused on small-diameter 

shafts that may be completely encased by a magnetic bearing housing. For the large-diameter shaft in this 

study, it was considered that a passive rotor may be applicable with a minimum of four electric magnetic 

poles at 90° apart, as shown in figure 46. The stiffness of the large-diameter system and radial growth 

after spinup are among the technical challenges to overcome. For example, an exoskeletal concept as 

studied would have a calculated radial growth of 0.040 in. at the bearing on the turbine end of the spool. 

Radial growth of this magnitude would result in stability problems; consequently, a magnet pole 

positioning system would be required to maintain the appropriate clearances for the operation of the 

magnetic bearing system. This positioning system would require high-speed sensing and positioning. The 

passive magnetic laminate and its mounting hardware will require high structural integrity to resist the 

extremely high inertial forces. This would most likely drive an increase in the weights; however, this is 

not considered for the present study in the total weight compilation. Although not inconsequential, the 

weight resulting from the magnetic bearing power consumption was not considered in this study. 

 Similar to the foil bearings, a backup bearing system is needed for off-nominal loading events. These 

events would include gust and hard-landing loads as well as loss-of-blade imbalance loads. The 

approximate sizing of the backup bearing system was done using a 5-g loading and a three-blade 

imbalance force of 76 000 lb. For this study, we assumed that this backup system would add 75 lb for the 

short-duration, off-nominal events considered. The above weight of 75 lb is only a place holder and 

should be considered as a low-fidelity number. This backup system would prevent excessive excursion 

from a centered position if temperatures were tolerable. Unfortunately, the high heat load prevents a 

system of this nature from functioning properly and without advances in lubrication and materials, would 

most certainly result in catastrophic bearing failure. 

Bearing weights.⎯The support structure for the bearing system (discussed in the next section) 

requires nominally a 47-lb total structure weight for the rings and struts at the forward and aft bearing 

support locations. The computed weights for a magnetic and foil bearing system as applicable in an 

exoskeletal design are tabulated in table 20. These weights will be considered as minimum weights since 

the historical data needed for sizing large-diameter, noncontact bearings in a flight propulsion system are 

not available. 
NASA/TM—2005-213369 46 

TABLE 20.⎯EXOSKELETAL ENGINE BEARING WEIGHT SUMMARY 

[All weights are in pounds.] 

Bearing concept alternatives 

Magnetic bearing system Foil bearing system 

Bearing support structure 46.6 Bearing support structure 46.6 

E-coils 96.8 Auxiliary power unit 84 

Positioning motors 95 Compressor 42 

Electrical controllers 30 Manifolding 24 

Permanent magnets 54.3 

Backup bearing 75 

Total weight 397.7 Total weight 196.6 

Summary.⎯The increase in diameter for a rotating drum versus the conventional shaft bearings 

creates bearing speeds well beyond the current operational capabilities for state-of-the-art lubricated 

bearing systems. This has prompted the consideration of other bearing schemes, such as a foil or 

magnetic systems. Both systems appear to meet the requirements of the exoskeletal application although 

neither technology is currently ready for operation at this size. Current developments in the foil bearing 

technologies indicate it may take 20 years to achieve foil bearings for this diameter. The magnetic 

bearings appear to be too heavy for this application and would also face a lengthy technology 

development program if selected. Regardless of these drawbacks, both bearing types were considered in 

the final comparison with the existing AE3007 high-pressure-spool weight comparison. 

Bearing Housing Study 

Description.⎯The bearing housing structure shown in figure 47 and modeled in figure 48 provides 

support for the exoskeletal rotor bearings located at each end of the rotor assembly. The fore and aft 

housings were designed as a ring structure attached to the engine nacelle frame through eight support 

vanes. The bearing rings and support vanes are made of 0.5-in.-thick titanium for strength and stiffness. 

 
NASA/TM—2005-213369 47 

 Analysis.⎯Stress and stability analyses were performed to size the bearing housing structure and to 

estimate its weight. The design load condition assumes that three side-by-side compressor blades have 

been detached, which results in an imbalance load of 7600 lb at 16 400 rpm. A maximum 8000-lb engine 

thrust load and the 7600-lb imbalance load were applied simultaneously to the bearing housing FEM, and 

a static analysis solution was executed in MSC/NASTRAN. Results of the static analysis show a peak 

Von Mises stress of 34 klb/in.2

 in the vanes and a maximum deformation of 0.0612 in. This stress result 

is well below the titanium compressive yield strength of 110 klb/in.2

 Using the same model, a linear 

buckling analysis was also performed to check the stability of the bearing ring support struts. The results 

are shown in figure 49. In summary, the bearing support configuration would sustain the applied service 

loads with sufficient margin. The corresponding structural weight was calculated to be 46.6 lb for both 

the fore and aft bearing housings. 

Findings, Observations, and Enabling Technology Identification 

Findings 

 (1) The exoskeletal high-pressure spool with a foil bearing system provides no weight advantage over 

a conventional design. The exoskeletal high-pressure spool with a magnetic bearing system weighs 

significantly more than the conventional design. This weight takes into account the expected weight 

savings that could be realized if a conventional engine were to make use of ceramic materials and switch 

to integral bladed disks. In addition, the exoskeletal weights assume that the negative margins in the 

drum rotor and blades could be overcome without additional weight impact. A summary of these weights 

is given in table 21. These findings reveal that the conceptual exoskeletal rotor and stator can be lighter 

than their conventional counterparts, subject to the assumptions and conditions of this study. However, 

the integrated exoskeletal high-pressure spool system is as heavy as or is heavier than its conventional 

counterpart primarily because of the bearing system mass. 
NASA/TM—2005-213369 48 

TABLE 21.⎯WEIGHT COMPARISON 

[All weights are in pounds.] 

Conventional Exoskeletal with magnetic bearings Exoskeletal with foil bearings 

Rotor Rotor drum 142.6 Rotor drum 142.6 

 HPC 187.3 

 HPT 79.7 

Total 267.0 

Stator Stator 100.5 Stator 100.5 

 HPC 111.1 

 HPT 54.1 

Total 165.2 

Bearings Bearing system Bearing system 

 Shaft Housings 46.6 Housings 46.6 

 Bearings E-coils 96.8 Auxiliary power unit 84.0 

Total 59.6 Position motors 95.0 Compressor 42.0 

Weight savings Controllers 30.0 Manifolds 24.0 

 Material change −38.7 Permanent magnets 54.3 

 Bladed disk design −10.0 Backup bearings 75.0 

Total −48.6 Total 397.7 Total 196.6 

Total system weight 443.2 Total system weight 640.9 Total system weight 439.7 

 (2) A requirement for turbine blade cooling is indicated in the exoskeletal design of an AE3007-like 

high-pressure spool examined in this study. Differential temperatures at the blade-to-ring interface could 

introduce excessive blade stresses. Blade cooling is a conventional approach to avoiding such thermalinduced stresses, but the development of cooling strategies for ceramic blades (which today are used to 

avoid cooling) is beyond the scope of this study and is a significant technology challenge. 

 (3) Shell material choices are limited by high temperatures and the need for light weight (low 

rotational inertia). The rotating drum operational environment is very demanding such that stiff, strong, 

low-density materials are required to sustain the high inertial forces and, at the same time, high thermal 

loads in the vicinity of the combustor and turbine require materials that can survive in an elevated 

temperature environment. Unfortunately, the thermal-resistant materials are very dense, and the strong 

lightweight materials are not temperature resistant. Within the scope of this investigation, no solution 

was identified using currently available materials, which is an indication that high-temperature, 

lightweight shell structures appear to be a critical design and technology challenge should an exoskeletal 

engine development be undertaken. 

 (4) Compressor blade analysis revealed that the taller metallic compressor blades, if used, in the 

exoskeletal high-pressure spool would have been buckling critical. Negative margins were also obtained 

in shear for compressor blades. These findings are a basis for considering composite materials for taller 

blade rows in exoskeletal compressors. 

Observations 

 (1) An inherent characteristic of the exoskeletal design approach is keeping rotational inertia low, 

which is opposed to providing a large central flow area for potential noise reduction or alternative cycle 

integration. The exoskeletal approach places blades and their supporting structure at the outer diameter of 

the system. Throttle response, maneuvering loads, and containment are among many factors that drive 
NASA/TM—2005-213369 49 

rotating machinery masses down at larger diameters and/or reduce rotational speeds. Since reductions in 

rotational speed directly impact engine performance, exoskeletal engines are driven toward lower 

diameters (which restricts the core space available for noise reduction or other cycle implementation). 

 (2) The increased rotary inertia of an exoskeletal design will require higher power for starting and 

have a slower response to changes in throttle setting. 

 (3) The fabrication of a woven preform of exoskeletal bladed rings would be difficult because of the 

complexity of the yarn paths and the large ring diameters. The densification of complex structures with 

tight surface tolerances has not been demonstrated. CMC’s are also difficult to densify without 

obstructing cooling passages. Machining the passages after densification would weaken the structure by 

cutting fibers. Inserting refractory metal rods in the cooling passages during densification could only be 

done if the passages were straight. 

Enabling Technology Identification 

 (1) Large-diameter bearing capability for high-speed bearings, noncontact or contact types: Even 

under the best of assumptions, rotor and stator weight savings are offset by the weight of a noncontact 

bearing system such as foil bearings. The use of magnetic bearings may be less of a technology leap, but 

they increase the system weight beyond that of conventional engine systems. 

 (2) Manufacturing capability for a bladed-ring component: The capability of manufacturing CMC 

ring blades as an integrated component with precise geometry and uniform mechanical properties needs 

to be developed. Durability under life-cycle and environmental conditions (especially critical thermal 

conditions as noted in the Findings section) needs to be demonstrated before composite materials could 

be considered in a gas turbine engine. 

Conclusions 

 1. The spool weight (neglecting bearings) of an exoskeletal engine may be equal to or less than the 

weight of the spool of a conventional engine. It was found that mounting blades inside a shell, rather than 

on a shaft, can result in rotating structure weights as low as or potentially lower than blades on shaftmounted disks. This conclusion is based on an aggressive preliminary assessment of a potential 

exoskeletal design approach. 

 2. An advancement in bearing technology is required before a more precise assessment of exoskeletal 

engine feasibility can be undertaken. Current bearing technologies are inadequate for the high rotational 

speeds at large diameters required in exoskeletal systems. The present maturity of enabling bearing 

technologies is insufficient to provide a precise assessment of the system implications of their integration 

in an exoskeletal engine. 

Recommendations 

 1. The next logical step in an examination of exoskeletal engine feasibility is to consider a “clean 

sheet” design approach to a specific exoskeletal design goal. The design of an exoskeletal system to its 

best advantage (especially involving changes to the engine cycle) was not attempted during this 

investigation. A clean sheet exoskeletal design, free of constraints driven by comparability to an existing 

engine, may discover alternatives that improve the weight and performance or lessen the technology 

challenges of the exoskeletal approach. Other applications of the exoskeletal concept have been 

suggested and may serve as compelling design goals. Noise reduction has been proposed through an 

inversion of the velocity profile at the convergence of “free flow” central core flow with the exhausts of 
NASA/TM—2005-213369 50 

the spools. The open central core has also been proposed for the integration of a ramjet or other cycle. An 

extension of the exoskeletal approach appears in the concept of vaneless counterrotating gas turbines. A 

clean sheet exoskeletal implementation for a specific design goal that examines competitive weight 

characteristics is a logical next step in the assessment of exoskeletal viability. 

 2. Exoskeletal implementation has been shown to be reliant on advancements in bearing technology. 

Therefore, a critical path to exoskeletal engine consideration is through bearing technology advancement. 

It is recommended that consideration of exoskeletal requirements be added to goals for high-diameterrotation-speed aerospace bearing technology programs. 
NASA/TM—2005-213369 51 

Appendix A

Weight Analysis Turbine Engine (WATE) Output for AE3007 

NEPP-WATE-98, March 1998 

 50 PAX ENGINE - Current Technology Baseline (revised for ' 

 &D NMODES=1,DRAW=T,LONG=F,DOUTHD=T,NCODE=-1,AMAC=F,NCASE=1,BOAT=F,ICEC=0, 

 TABLES=T,PINPUT=T,TLOAD=F,INST=0,MAPLOT=F,MAXNIT=100,ITERM=1,IWT=1, &END 

 &D MODE=1,CALBLD=F,CBFLOW=0.020,ERCB=0.80,ICBCMP=5, 

 KONFIG(1,1)=1,1,0,2,0,SPEC(1,1)=260.0,0,0,0,0.00,8600,1,1,0.,0,4.44,27., 

 KONFIG(1,2)=4,2,0,3,0,SPEC(1,2)=1.60,0,1,3401,1,3402,1,3403,1,0, 

 0,.8750,1.62,0.9000, 

 KONFIG(1,3)=7,3,0,4,20,SPEC(1,3)=5.3,0.00,0.015, 

 KONFIG(1,4)=2,4,0,5,0,SPEC(1,4)=.010, 

 KONFIG(1,5)=4,5,0,6,21,SPEC(1,5)=1.25,0.115,1,3707,1,3708,1,3709,1,0, 

 .0,.8600,15.0,0.950, 

 KONFIG(1,6)=2,6,0,7,0,SPEC(1,6)=.015,8*0, 

 KONFIG(1,7)=2,7,0,8,0,SPEC(1,7)=.045,0.,0.,2760.,0.995,18400,5*0.,6, 

 KONFIG(1,8)=5,8,21,9,0,SPEC(1,8)=3.45,0.977,1,3801,1,3802,1,1, 

 .6850,1,.9100,7000.,1,1122,2, 

 KONFIG(1,9)=2,9,0,10,0,SPEC(1,9)=0.005,6*0.,0.55, 

 KONFIG(1,10)=5,10,0,15,0,SPEC(1,10)=2.10,0.,1,3803,1,3804,1,1, 

 1.00,1,.920,5000.0,1,8888,4, 

 KONFIG(1,14)=2,20,0,22,0,SPEC(1,14)=0, 

 KONFIG(1,15)=2,15,0,16,0,SPEC(1,15)=0.01, 

 KONFIG(1,11)=8,16,22,12,0,SPEC(1,11)=0,0,0.3,.95,1, 

 KONFIG(1,12)=2,12,0,13,0,SPEC(1,12)=.0065,8*0., 

KONFIG(1,13)=9,13,0,14,0,SPEC(1,13)=0.0,7004,0.,0.,7003,0.,0.,0.,1.,3*0,0.5,11

, 

 KONFIG(1,21)=11,5,8,23,0,SPEC(1,21)=18000.,8*1, 

 KONFIG(1,22)=11,2,10,0,0,SPEC(1,22)=7000.,0.435,0.435,6*1, 

 KONFIG(1,23)=10,0,0,0,0,SPEC(1,23)=-100., 

 

 /* Design Point CNTLs... IF NEEDED */ 

 KONFIG(1,31)=12,SPCNTL(1,31)=1,1,'STAP',1,5,35.5,0, 

 KONFIG(1,32)=12,SPCNTL(1,32)=1,5,'DOUT',5,5,15.0,0, 

 KONFIG(1,33)=12,SPCNTL(1,33)=13,2,'STAP',2,6,345.7,0, 

 KONFIG(1,34)=12,SPCNTL(1,34)=4,7,'PERF',4,0,7580.,1, 

 KONFIG(1,35)=12,SPCNTL(1,35)=4,7,'PERF',4,0,1955.,0, 

 

 /* Off-Design CNTLs */ 

 KONFIG(1,41)=12,0,0,0,0,SPCNTL(1,41)=1,1,'STAP',8,13,0,0,0,0, 

 KONFIG(1,42)=12,0,0,0,0,SPCNTL(1,42)=1,2,'STAP',8,2,0,0,0,0, 

 KONFIG(1,43)=12,0,0,0,0,SPCNTL(1,43)=1,5,'STAP',8,5,0,0,0,0, 

 KONFIG(1,44)=12,0,0,0,0,SPCNTL(1,44)=1,8,'STAP',8,8,0,0,0,0, 

 KONFIG(1,45)=12,0,0,0,0,SPCNTL(1,45)=1,10,'STAP',8,10,0,0,0,0, 

 KONFIG(1,46)=12,0,0,0,0,SPCNTL(1,46)=2,11,'DOUT',8,11,0,0,0,0, 

 KONFIG(1,50)=12,0,0,0,0,SPCNTL(1,50)=1,21,'DOUT',8,21,0,0,0,0, 

 KONFIG(1,51)=12,0,0,0,0,SPCNTL(1,51)=1,22,'DOUT',8,22,0,0,0,0, 

 

 /* Part Power and Operational Controls */ 

 KONFIG(1,55)=16,0,0,0,0,SPEC(1,55)=1,57,'DOUT',4,2,1.1,0,1,-1,0,1,9, 

 KONFIG(1,56)=16,0,0,0,0,SPEC(1,56)=0,58,'DOUT',4,2,1.1,0,1,-1,0,1,9, 

 KONFIG(1,57)=12,0,0,0,0,SPCNTL(1,57)=1,3,'DOUT',5,2,23.0,0,0,0, 

 KONFIG(1,58)=12,0,0,0,0,SPCNTL(1,58)=1,3,'STAP',6,22,0.70,0,0,0, 

 

 /* control for marching via temp. */ 

 KONFIG(1,60)=12,SPCNTL(1,60)=4,7,'PERF',4,0,7580,0,0, 

 &END 
NASA/TM—2005-213369 52 

 JFLOW = (FLOCAL EXECUTION ORDER) 

 1 2 3 4 5 6 7 8 9 10 15 14 11 

12 13 0 0 0 0 0 

 1 

 < INLT 1> 

 2 

 < COMP 2> 

 3 

 < SPLT 3> < SPLT 3> 

 4 20 

 < DUCT 4> < DUCT 14> 

 5 22 

 < COMP 5> < COMP 5> < MIXR 11> 

 21 6 

 < TURB 8> < DUCT 6> 

 7 

 < BRNR 7> 

 8 

 < TURB 8> 

 9 

 < DUCT 9> 

 10 

 < TURB 10> 

 15 

 < DUCT 15> 

 16 

 < MIXR 11> 

 12 

 < DUCT 12> 

 13 

 < NOZZ 13> 

 14 

 

0SHAFT (21) IS CONNECTED TO ( 5) AND ( 8) AND (23) AND 

0SHAFT (22) IS CONNECTED TO ( 2) AND (10) AND 

1 

0 THE FOLLOWING REPRESENTS THE CONFIGURATION FOR MODE= 1 

 50 PAX ENGINE - Current Technology Baseline (revised for ' 

 CONFIGURATION DATA 22 STATIONS 60 COMPONENTS 

 COMPONENT NKIND COMPONENT UPSTREAM DOWNSTREAM 

 NUMBER TYPE STATIONS STATIONS 

 1 1 INLT 1 0 2 0 

 2 4 COMP 2 0 3 0 

 3 7 SPLT 3 0 4 20 

 4 2 DUCT 4 0 5 0 

 5 4 COMP 5 0 6 21 

 6 2 DUCT 6 0 7 0 

 7 2 BRNR 7 0 8 0 

 8 5 TURB 8 21 9 0 

 9 2 DUCT 9 0 10 0 

 10 5 TURB 10 0 15 0 

 11 8 MIXR 16 22 12 0 

 12 2 DUCT 12 0 13 0 

 13 9 NOZZ 13 0 14 0 

 14 2 DUCT 20 0 22 0 
NASA/TM—2005-213369 53 

 15 2 DUCT 15 0 16 0 

 21 11 SHFT 5 8 23 0 

 22 11 SHFT 2 10 0 0 

 23 10 LOAD 0 0 0 0 

 VARIABLE CONTROL INFORMATION 

 ACTIVE: 

 CNTL 34: vary des. exit T or f/a of BRNR 7 

 so that net jet thrust = 0.75800D+04 

 INACTIVE: 

 CNTL 31: vary weight-flow rate of 1 

 so that weight-flow rate at station 5 = 0.35500D+02 

 CNTL 32: vary "R" value of 5 

 so that % surge margin of 5 = 0.15000D+02 

 CNTL 33: vary design PR of 2 

 so that total pressure at station 6 = 0.34570D+03 

 CNTL 35: vary des. exit T or f/a of BRNR 7 

 so that net jet thrust = 0.19550D+04 

 CNTL 41: vary weight-flow rate of 1 

 so that flow-rate error at station 13 = 0.00000D+00 

 CNTL 42: vary "R" value of 2 

 so that flow-rate error at station 2 = 0.00000D+00 

 CNTL 43: vary "R" value of 5 

 so that flow-rate error at station 5 = 0.00000D+00 

 CNTL 44: vary PR used to read map of 8 

 so that flow-rate error at station 8 = 0.00000D+00 

 CNTL 45: vary PR used to read map of 10 

 so that flow-rate error at station 10 = 0.00000D+00 

 CNTL 46: vary inlet area JM2 flow of 11 

 so that static press. error of 11 = 0.00000D+00 

 CNTL 50: vary RPM or speed ratio of 21 

 so that (HP)net/(HP)total of 21 = 0.00000D+00 

 CNTL 51: vary RPM or speed ratio of 22 

 so that (HP)net/(HP)total of 22 = 0.00000D+00 

 CNTL 57: vary bypass flow ratio of 3 

 so that % surge margin of 2 = 0.23000D+02 

 CNTL 58: vary bypass flow ratio of 3 

 so that Mach number at station 22 = 0.70000D+00 

 CNTL 60: vary des. exit T or f/a of BRNR 7 

 so that net jet thrust = 0.75800D+04 

 CONDITIONAL CONTROL INFORMATION 

 ACTIVE: 

 VCNT 55: watch "R" value for maps of 2; 

 trigger value is 0.11000D+01 which turns on/off switch of 

CNTL 57 to 0.10000D+01 

 VCNT 56: watch "R" value for maps of 2; 

 trigger value is 0.11000D+01 which turns on/off switch of 

CNTL 58 to 0.00000D+00 

 INACTIVE: 

 None!!! 
NASA/TM—2005-213369 54 

1 

 INPUT - Warning: the maximum component number used, 60, does not equal the 

number of components configured in any one mode, 36! 

 UPDATED INPUT DATA TO REFLECT CALCULATED INPUT 

 COMPONENT 

 NO. TYPE SPEC 1 SPEC 2 SPEC 3 SPEC 4 SPEC 5 

SPEC 6 SPEC 7 SPEC 8 SPEC 9 

 1 INLT 260.0000000 0.000000000 14.69600000 0.000000000 0.000000000 

8600.000000 1.000000000 1.000000000 0.000000000 

 2 COMP 1.600000000 0.000000000 3298.567025 3401.000000 0.871762004 

3402.000000 0.974387528 3403.000000 1.561712846 

 3 SPLT 5.300000000 0.000000000 0.015000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 4 DUCT 0.010000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 5 COMP 1.250000000 0.115000000 17088.15477 3707.000000 29.85406452 

3708.000000 0.869525830 3709.000000 2.237728006 

 6 DUCT 0.015000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 7 BRNR 0.045000000 0.000000000 0.000000000 2792.143036 0.995000000 

18400.00000 0.000000000 0.000000000 0.000000000 

 8 TURB 3.450000000 0.977000000 1.108284344 3801.000000 0.196672238 

3802.000000 1.028548856 1.460070201 0.685000000 

 9 DUCT 0.005000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.550000000 0.000000000 

 10 TURB 2.100000000 0.000000000 0.314600153 3803.000000 0.259514065 

3804.000000 1.009740547 2.554108500 1.000000000 

 11 NOZZ 327.6979091 508.8929531 0.300000000 0.950000000 1.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 12 DUCT 0.006500000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 13 NOZZ 696.0257350 7004.000000 0.000000000 0.000000000 7003.000000 

0.000000000 0.000000000 1.000000000 1.000000000 

 14 DUCT 0.000000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 15 DUCT 0.010000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 21 SHFT 18000.00000 1.000000000 1.000000000 1.000000000 1.000000000 

1.000000000 1.000000000 1.000000000 1.000000000 

 22 SHFT 7000.000000 0.435000000 0.435000000 1.000000000 1.000000000 

1.000000000 1.000000000 1.000000000 1.000000000 

 23 LOAD -100.0000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

0MODE 1 NOW BEING USED 

0CASE IDENTIFICATION 50 PAX ENGINE - Current Technology Baseline (revised 

for ' 

 STATION PROPERTY OUTPUT DATA 

 FLOW WEIGHT TOTAL TOTAL FUEL/AIR CORRECTED 

MACH STATIC INTERFACE CORRECTED 

 STATION FLOW PRESSURE TEMPERATURE RATIO FLOW 

NUMBER PRESSURE FLOW ERROR 

 NNEP-ARP? STATP1 STATP2 STATP3 STATP4 STATP5 

STATP6 STATP7 STATP8 

 1-1 257.4356705 14.69600000 545.6700000 0.000000000 264.0504947 

0.000000000 0.000000000 0.000000000 
NASA/TM—2005-213369 55 

 2-2 257.4388394 14.54904000 545.6700000 0.000000000 266.7176714 

0.000000000 0.000000000 -0.000012309 

 3-3 257.4388394 23.97330869 641.9425052 0.000000000 175.5684700 

0.000000000 0.000000000 0.000000000 

 4-4 41.68325273 23.97330869 641.9425052 0.000000000 28.42719819 

0.000000000 0.000000000 0.000000000 

 5-5 41.67950672 23.73357560 641.9425052 0.000000000 28.71434160 

0.000000000 0.000000000 0.000089872 

 6-6 36.05277331 353.4369834 1465.037705 0.000000000 2.519436716 

0.000000000 0.000000000 0.000000000 

 7-7 36.05277331 348.1354286 1465.037705 0.000000000 2.557803773 

0.000000000 0.000000000 0.000000000 

 8-8 36.84889128 332.4693343 2792.143036 0.022090407 3.779180147 

0.000000000 0.000000000 0.82086D-05 

 9-9 41.53179226 72.39087045 1942.626873 0.019550915 16.31713615 

0.000000000 0.000000000 0.000000000 

 10-10 41.53188612 72.02891609 1942.626873 0.019550915 16.39913181 

0.000000000 0.000000000 -0.22600D-05 

 12-12 257.2874728 21.04741771 775.7310774 0.003105043 219.6988497 

0.519548746 17.52962411 0.000000000 

 13-13 257.2470341 20.91060949 775.7310774 0.003105043 221.1362353 

0.723876953 14.69600000 0.000312400 

 14-14 257.2470341 20.91060949 775.7310774 0.003105043 221.1014786 

0.721975555 14.69600000 0.000000000 

 15-15 41.53188612 18.14278133 1426.019774 0.019550915 55.78185202 

0.000000000 0.000000000 0.000000000 

 16-16 41.53188612 17.96135352 1426.019774 0.019550915 56.34530507 

0.282217026 17.03306232 0.000000000 

 20-20 215.7555866 23.61370906 641.9425052 0.000000000 149.3820019 

0.000000000 0.000000000 0.000000000 

 21-21 4.793143273 353.4369834 1465.037705 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 

 22-22 215.7555866 23.61370906 641.9425052 0.000000000 149.3820019 

0.699577332 17.04007626 0.000000000 

 COMPONENT OUTPUT DATA 

 COMPONENT 

 NO. TYPE DATOUT1 DATOUT2 DATOUT3 DATOUT4 DATOUT5 

DATOUT6 DATOUT7 DATOUT8 DATOUT9 

 Ram drag V-fps V-knots T2/T1 P2/P1 

Mach (Eta)r T2/518.67 Alt-ft 

 1 INLT 0.000000000 0.000000000 0.000000000 1.000000000 1.000000000 

0.000000000 0.990000000 1.052056221 0.000000000 

 HP in RPM or ratio Z-3Dmap R-map % Margin 

Corr. Spd. (Eta)poly. (Eta)adia. (PR)comp. 

 2 COMP -8427.558152 3073.102608 0.000000000 1.516729182 19.54189788 

0.908306189 0.876769658 0.867798381 1.647758800 

 Bypass (dP/P)main (dP/P)2nd Tot W main React. W 

main 

 3 SPLT 5.176073663 0.000000000 0.015000000 257.4388394 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 

 4 DUCT 0.000000000 0.010000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 641.9425052 

 HP in RPM or ratio Z-3Dmap R-map % Margin 

Corr. Spd. (Eta)poly. (Eta)adia. (PR)comp. 

 5 COMP -12123.73271 18006.80836 0.000000000 1.251725672 28.99598420 

0.947194291 0.902213785 0.860591837 14.89185571 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 
NASA/TM—2005-213369 56 

 6 DUCT 0.000000000 0.015000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 1465.037705 

 (dP/P)mom. dP/P Des. M in Fuel/Wa A-sq. in. 

Fuel Flow Mach In (Eta)burn. T burn. 

 7 BRNR 0.000000000 0.045000000 0.000000000 0.022090407 0.000000000 

2867.113603 0.000000000 0.995000000 2792.143036 

 HP out RPM or ratio Z-3Dmap PR-Table (Eta)poly. 

Corr. Spd. TRIT,deg R (Eta)adia. (PR)turb. 

 8 TURB 12223.74322 18006.80836 1.000000000 3.460633063 0.894445188 

7002.647694 2693.665432 0.909987829 4.592697011 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 

 9 DUCT 0.000000000 0.005000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 1942.626873 

 HP out RPM or ratio Z-3Dmap PR-Table (Eta)poly. 

Corr. Spd. TRIT,deg R (Eta)adia. (PR)turb. 

 10 TURB 8428.951686 3073.102608 1.000000000 2.162877116 0.907538139 

5047.414844 1942.626873 0.920815070 3.970114327 

 Amain-sq in A2nd-sq in Mom. Coef. Effectiv. V-main 

V-2nd V-exit (Del P)s Aexit/Aent 

 11 NOZZ 360.1563202 476.4345421 0.950000000 0.500000000 507.9996455 

829.1265337 689.4721929 0.000411699 1.000000000 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 

 12 DUCT 0.000000000 0.006500000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 775.7310774 

 F gross VJ act Pt/Ps Aex-sq in Ath-sq in 

(CD)flow (CV)vel FG Coeff. (PR)nozz 

 13 NOZZ 7467.934584 942.6401219 1.422877619 689.6592463 689.6592463 

0.996935621 0.997373314 0.990853084 1.422877619 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 

 14 DUCT 0.000000000 0.000000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 641.9425052 

 (dP/P)mom. dP/P Des. M in A-sq. in. 

Mach in T exit 

 15 DUCT 0.000000000 0.010000000 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 1426.019774 

 Net HP RPM or ratio RPM comp. 1 RPM comp. 2 RPM comp. 3 

RPM comp. 4 HP net/HP tot 

 21 SHFT 0.010503844 18006.80836 18006.80836 18006.80836 18006.80836 

0.000000000 0.000000000 0.85930D-06 0.000000000 

 Net HP RPM or ratio RPM comp. 1 RPM comp. 2 RPM comp. 3 

RPM comp. 4 HP net/HP tot 

 22 SHFT 1.393533823 7064.603696 3073.102608 3073.102608 0.000000000 

0.000000000 0.000000000 0.000165341 0.000000000 

 Load HP RPM or ratio 

 23 LOAD -100.0000000 18006.80836 0.000000000 0.000000000 0.000000000 

0.000000000 0.000000000 0.000000000 0.000000000 

 CONTROL INFORMATION 

 ACTIVE CNTLs: 46 45 44 43 58 42 41 50 51 

 ACTIVE VNTLs: 55 56 

 MACH= 0.0000 ALTITUDE= 0. RECOVERY= 0.9900 

 4 ITERATIONS 14 PASSES 

 AIRFLOW (LB/SEC) 257.44 GROSS THRUST 7467.93 

FUEL FLOW (LB/HR) 2867.11 
NASA/TM—2005-213369 57 

 NET THRUST 7467.93 TSFC 0.3839 

NET THRUST/AIRFLOW 29.0089 

 INLET RAM/ADDED DRAG 0.00 TOTAL BRAKE SHAFT HP 1.40 

BOATTAIL DRAG 0.00 

 INSTALLED THRUST 7467.93 INSTALLED TSFC 0.3839 

SPILLAGE + LIP DRAG 0.00 

 TOTAL NACELLE DRAG 0.00 CET (T4) 2792.1 

RIT (T41) 2693.7 

 CORR AIRFLOW (LB/SEC) 266.72 

 FNIN1 7467.93 SFCIN1 0.3839 

 FNIN2 7467.93 SFCIN2 0.3839 

 EI (G NOX/KG FUEL) 28.555 

 MASS AVG. VJET 942.64 

 PROPULSIVE EFFICIENCY 0.0000000 

0 CUSTOMER BLEED WAS EXTRACTED FROM COMPRESSOR 5 

 TOTAL TEMPERATURE = 1306.6705 TOTAL PRESSURE = 236.2536 WEIGHT FLOW = 

0.8336 ENERGY RATIO = 0.8000 

NASA/TM—2005-213369 59 

Appendix B

NASA Engine Performance Program (NEPP) Output 

&D IWT=2, &END 

0MODE 1 NOW BEING USED 

 &W 

 IPLT = T, 

 ISII = F, 

 ISIO = F, 

 PLOT = T, 

 SKIPIT = F, 

 ILENG = 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 

 11, 12,13, 

 ACCS = 0.0950, 

 DISKWC = 1., 

 DISKWI = 1., 

 DISKWT = 1., 

 

 IWMEC(1, 1) = 1, 12, 1, 10, -12, 

 1, 0, 0, 0, 1, 

 0, 

 IWMEC(1, 2) = 48, 3, 0, 1, 1, 

 0, 0, -2, 0, 1, 

 0, 0, 0, 0, 0, 

 IWMEC(1, 3) = 7, 0, 

 IWMEC(1, 4) = 2, 2, 0, -1, 

 IWMEC(1, 5) = 47, 3, 0, 1, 14, 

 -3, 0, 0, 12, 0, 

 2, 0, 0, 0, 0, 

 IWMEC(1, 6) = 2, 2, 0, 

 IWMEC(1, 7) = 21, 1, 1, 

 IWMEC(1, 8) = 51, 3, 0, 0, 2, 

 0, 5, 0, 0, 5, 

 0, 0, 1, 0, 1, 

 IWMEC(1, 9) = 2, 2, 0, 

 IWMEC(1, 10) = 52, 3, 0, 0, 3, 

 1, -6, 0, 0, 2, 

 0, 0, 0, 0, 1, 

 IWMEC(1, 11) = 82, 0, 

 IWMEC(1, 12) = 2, 1, 0, 

 IWMEC(1, 13) = 9, 1, 2, 0, 1, 

 1, 5, 0, 0, 0, 

 IWMEC(1,14) = 2, 3, 0, -4, 

 IWMEC(1,15) = 2, 1, 

 IWMEC(1, 21) = 11, 2, 8, 0, 0, 

 0, 5, 2, 

 IWMEC(1, 22) = 11, 1, 10, 0, 0, 

 0, 2, 3, 

 FRVAL(1, 1) = 0.11, 1., 0.1, 1.0, 0, 

 0.10, 0.1, 0, 0, 0, 

 0.1, 0, 1, 2, 2, 

 0, 0, 0, 

 FRVAL(1, 2) = 0.12, 1.0, 0.1, 1.4, 6, 

 0.12, 0.2, 0, 0, 0, 

 1., 0, -5, 0, 0, 

 1, 100., 0.9, 
NASA/TM—2005-213369 60 

 FRVAL(1, 3) = 0.12, 1.0, 0.1, 0.8, 0, 

 0.12, 0.2, 0, 0, 0, 

 0.00, 1, -2, 2, 3, 

 0, 0, 0, 

 FRVAL(1, 4) = 0.12, 1.0, 0.1, 1.8, 3, 

 0.12, 0.1, 0, 0, 0, 

 5., 0, 0, 0, 0, 

 0, 0, 0, 

 FRVAL(1, 6) = 0.12, 1.0, 0.1, 1.2, 6, 

 0.12, 0.2, 0, 0, 0, 

 0.50, 0, 0, 5, 5, 

 0, 0, 0, 

 DESVAL(1, 1) = .80, 1000., 0., 0., 0, 

 1, 1, 1, 0, 0, 

 0.4, 0, 1650., 70., 0, 

 0.1, 0.1, 0.09, 0, 0, 

 8, 

 DESVAL(1, 2) = .550, .407, 0, 0, 1.8, 

 1, 0, 0.39, 0, 0, 

 0, 0, 0, 0, 0, 

 24, 0, 1.1, 0.024, 0.10, 

 2.8, 2.8, 1.4, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 50, 0, 0, 0.135, 5.9, 

 1., 1., 2.5, 2.5, 0, 

 0, 1, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 1., 0.6, 1.5, 1.1, 0.05, 

 0, 0, 0, 0, 0, 

 DESVAL(83,2) = 1360., 

 DESVAL(1, 3) = 

 DESVAL(1, 4) = .500, 0, 0, 0, 0, 

 5.0, -1, 0, 0, 

 DESVAL(1, 5) = 0.34, 0.31, 0, 0, 1.4, 

 1, 1.0, 0.455, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0.84, 0.162, 0.08, 

 2.2, 0.8, 0.833, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0.75, 0.163, 0.18, 

 1.0, 1.0, 3.1, 1.2, 0, 

 0, 5, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 36, 0, 0, 2.5, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 1000., 0, 9, 

 DESVAL(1, 6) = 0.3, 0, 0, 0, 0, 

 4., -1, 0, 0, 0, 

 DESVAL(1, 7) = 50, 0.013, 0, 0, 

 0, 0, 

 DESVAL(16,7) = 0.29, 0.29, 

 DESVAL(1, 8) = .100, .300, 0.34, 0, 0, 

 1, 1, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, .281, 0.71, 0.70, 0.2, 

 1.2, 1.8, 1., 
NASA/TM—2005-213369 61 

 DESVAL(36, 8)= 0, 0.28, 0.92, 0.38, 0.17, 

 DESVAL(76, 8)= 0.28, 0.32, 100000., 

 DESVAL(1, 9)= 0.3, 0, 0, 0, 0, 

 0.8, -1, 0, 0, 

 DESVAL(1, 10)= .165, .319, .280, 0, 0, 

 1.3, 1., 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, .286, 1.22, 0.20, 0.2, 

 2.8, 3.7, 1., 

 DESVAL(36,10)= 0, 0.28, 1.11, 0.057, 0.12, 

 DESVAL(76,10)= 0.28, 0.32, 100000, 

 DESVAL(1, 11)= 0.64, 12, 1.0, 0, 0, 

 0, 0, 0, 0, 0, 

 0, 0, 0, 0, 0, 

 0.156, 0.1, 

 DESVAL(1, 12)= 0.41, 0, 0, 0, 0, 

 0.25, -1, 0, 0, 

 DESVAL(1, 13)= 0.18, 1000., 0, 0.5, 8.3, 

 0, 0, 0, 0, 0, 

 1, 0, 0.8, 0, 0, 

 0.281, 0.1, 1., 1., 0, 

 DESVAL(26,13)= 0, 0.1, 1., 1., 

 DESVAL(1, 14) = 0.45, 0, 0, 0, 0, 

 0, -1, 0, 

 DESVAL(7, 21)= 80000, 0.6, 5, 8, 0, 

 0, 0, 0, 0, 0.30, 

 DESVAL(21, 21)= 0.186, 3, 21., 0.5, 

 DESVAL(26, 21)= 0.186, -4, 60., 0.5, 

 DESVAL(7, 22)= 80000, 0.6, 2, 10, 0, 

 0, 0, 0, 0, 0.30, 

 DESVAL(21, 22)= 0.186, 1, 5., 0.5, 

 DESVAL(26, 22)= 0.186, 2, 6.5, 0.5, 

 DESVAL(31, 22)= 0.186, -5, 70., 0.5, 

 &END 

 

 WEIGHT INPUT IS IN ENGLISH UNITS 

 WEIGHT OUTPUT IS IN ENGLISH UNITS 

 

 INPUT - Warning: the maximum component number used, 22, does not equal the 

number of components configured in any one mode, 17! 

 

 ************* 

 * * 

 * FANH 2 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 **************************************** 

 ALT MN VALUE 

 PTOT 0. 0.000 14.5 LB/SQIN 

 TTOT 0. 0.000 545.7 DEG R 

 CWIN 35000. 0.700 316.3 LB/SEC 

 **************************************** 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.550 612. 546. 2095. 1706. 6.8408 1.4001 

 UTIPMAX STRESS DEN W/AREA TR H/T 

 1519.9 71855.2 0.160 2.127 1.400 0.390 
NASA/TM—2005-213369 62 

 COMPRESSOR 2 MECHANICAL DESIGN 

 LOADING N STG DIAM U TIP C RPM C RPM MAX RPM 

 1.031 1.00 38.46 1360.0 8312.4 8104.1 9056.7 

 STAGE 1 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 52. 52. 36. 6. 0. 43. 40. 9. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 5.9 14.9 0.160 0.160 0.160 0.160 0.050 2.80 2.50 50 1.10 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.6198 22.1 0.550 6.841 7.50 19.23 24 1519.9 71855. 238. 546. 546. 

6724. 

 COMPRESSOR COMPONENT WEIGHT SUMMARY 

 WTDSK WTBLD WTDRM WTST/IGV WTSF WTCAS WTNAB WTCR 

 52.1 51.5 9.0 35.6 6.1 43.3 0.0 40.5 

 FRAME COMPONENTS INFORMATION 

 FRAME WT LENGTH GAP #FRAMES ARF RHOF 

 12.57 5.0 0.0 6 1.40 0.120 

 CASE WT HUB WT UP SUPP LW SUPP RHOS THSUP WTTOWER TS DIAM 

 14.49 9.22 0.00 0.00 0.120 0.20 19.89 1.00 

 FRAME WT = 56.17 

 N STG WEIGHT LENGTH CENGRA INERTIA LENGTH2 

 1 294.37 5.92 15.9 6724.1 19.94 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.407 495. 638. 3394. 3029. 5.7070 1.3981 

 PR AD EF PO TO HP 

 1.6200 0.8750 3394.0 637.7 8132. 

 HI HO WI CWI 

 130.43 152.54 260.00 269.37 

 ******************* TOTAL COMP WEIGHT IS 294.366 

 

 ************* 

 * * 

 * DUCT 4 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 FRAME COMPONENTS INFORMATION 

 FRAME WT LENGTH GAP #FRAMES ARF RHOF 

 2.40 1.7 0.2 45 1.00 0.110 
NASA/TM—2005-213369 63 

 CASE WT HUB WT UP SUPP LW SUPP RHOS THSUP WTTOWER TS DIAM 

 1.30 1.12 10.77 6.85 0.100 0.10 0.00 1.00 

 FRAME WT = 22.43 

 DUCT , 4 

 RH= 10.51 RT= 12.20 LENG= 8.46 

 AREA= 0.838 RHO=.168 

 WTC(OUTER) WTC(INNER) WTT(REV) WT(TOTAL) TMIN 

 5.4467 4.6912 0.0000 32.5692 0.0500 

 

 ************* 

 * * 

 * HPC 5 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 **************************************** 

 ALT MN VALUE 

 PTOT 0. 0.000 23.7 LB/SQIN 

 TTOT 0. 0.000 641.9 DEG R 

 CWIN 35000. 0.700 31.4 LB/SEC 

 **************************************** 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.340 416. 638. 3360. 3102. 1.0638 1.3981 

 UTIPMAX STRESS DEN W/AREA TR H/T 

 1121.4 50562.4 0.300 1.211 0.833 0.455 

 COMPRESSOR 5 MECHANICAL DESIGN 

 LOADING N STG DIAM U TIP C RPM C RPM MAX RPM 

 0.974 14.00 15.68 1000.0 16202.9 14613.0 16387.3 

 FRAME COMPONENTS INFORMATION 

 FRAME WT LENGTH GAP #FRAMES ARF RHOF 

 13.17 5.3 0.0 9 0.80 0.120 

 CASE WT HUB WT UP SUPP LW SUPP RHOS THSUP WTTOWER TS DIAM 

 6.32 2.87 13.36 3.97 0.120 0.20 0.00 1.00 

 FRAME WT = 39.70 

 STAGE 1 

 VARIABLE INLET GUIDE VANE PRESENT 

 WIGV RHOIGV ARIGV NBIGV 

 8. 0.160 2.50 36 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 4. 3. 2. 2. 8. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 4.5 4.5 0.160 0.160 0.160 0.160 0.000 2.20 3.10 26 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 
NASA/TM—2005-213369 64 

 1.3363 14.7 0.340 1.064 3.57 7.84 21 1121.4 26967. 33. 638. 642. 

107. 

 STAGE 2 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 7. 2. 2. 1. 1. 6. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 3.4 3.4 0.160 0.160 0.160 0.160 0.000 2.09 2.95 35 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.3043 14.7 0.338 0.838 4.80 7.84 28 1121.4 21341. 20. 699. 701. 

127. 

 STAGE 3 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 1. 1. 1. 1. 5. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 2.7 2.7 0.160 0.160 0.160 0.160 0.000 1.98 2.81 44 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.2779 14.7 0.336 0.674 5.53 7.84 35 1121.4 17202. 17. 759. 760. 

137. 

 STAGE 4 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 1. 1. 1. 1. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 2.3 2.3 0.160 0.160 0.160 0.160 0.000 1.88 2.66 53 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.2558 14.7 0.334 0.552 6.02 7.84 42 1121.4 14095. 15. 820. 818. 

142. 

 STAGE 5 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 1. 1. 1. 1. 3. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.9 1.9 0.160 0.160 0.160 0.160 0.000 1.77 2.52 62 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.2372 14.7 0.331 0.458 6.36 7.84 49 1121.4 11716. 13. 880. 877. 

146. 

 STAGE 6 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 0. 0. 1. 1. 3. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.7 1.7 0.160 0.160 0.160 0.160 0.000 1.66 2.37 71 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.2210 14.7 0.329 0.386 6.62 7.84 56 1121.4 9859. 11. 940. 936. 

146. 

 STAGE 7 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 6. 0. 0. 1. 1. 3. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.5 1.5 0.160 0.160 0.160 0.160 0.000 1.55 2.22 80 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 
NASA/TM—2005-213369 65 

 1.2071 14.7 0.327 0.328 6.82 7.84 62 1121.4 8387. 11. 999. 995. 

146. 

 STAGE 8 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 5. 0. 0. 0. 1. 2. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.4 1.4 0.160 0.160 0.160 0.160 0.000 1.45 2.08 88 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1949 14.7 0.325 0.281 6.97 7.84 68 1121.4 7204. 10. 1058. 1053. 

145. 

 STAGE 9 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 5. 0. 0. 0. 0. 2. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.3 1.3 0.160 0.160 0.160 0.160 0.000 1.34 1.93 95 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1841 14.7 0.323 0.244 7.09 7.84 74 1121.4 6239. 9. 1117. 1112. 

144. 

 STAGE 10 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 11. 0. 0. 0. 0. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.2 1.2 0.300 0.300 0.160 0.300 0.000 1.23 1.78 101 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1745 14.7 0.321 0.213 7.19 7.84 78 1121.4 10207. 17. 1175. 1171. 

305. 

 STAGE 11 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 11. 0. 0. 0. 0. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.2 1.2 0.300 0.300 0.160 0.300 0.000 1.12 1.64 106 3.00 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1659 14.7 0.319 0.187 7.28 7.84 82 1121.4 8967. 17. 1233. 1230. 

306. 

 STAGE 12 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 11. 0. 0. 0. 0. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.1 1.1 0.300 0.300 0.160 0.300 0.000 1.02 1.49 110 3.00 

 

   WARNING FOLLOWING STAGE DESIGN LIMIT EXCEEDED ** *** 

 STAGE HUBTIP RATIO IS 0.94 DES LIMIT IS 0.94 

 ** HUB TIP RATIO IS TOO HIGH REDUCE HUB TIP RATIO INPUT ** 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1582 14.7 0.316 0.165 7.34 7.84 84 1121.4 7926. 17. 1290. 1289. 

309. 

 STAGE 13 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 11. 0. 0. 0. 0. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 
NASA/TM—2005-213369 66 

 1.1 1.1 0.300 0.300 0.160 0.300 0.000 0.91 1.35 112 3.00 

 

   WARNING FOLLOWING STAGE DESIGN LIMIT EXCEEDED ** *** 

 STAGE HUBTIP RATIO IS 0.94 DES LIMIT IS 0.94 

 ** HUB TIP RATIO IS TOO HIGH REDUCE HUB TIP RATIO INPUT ** 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1512 14.7 0.314 0.147 7.40 7.84 85 1121.4 7044. 17. 1347. 1347. 

315. 

 STAGE 14 

 WD WB WS WSSF WN WC WFCR WTDRUM 

 11. 0. 0. 0. 0. 4. 0. 1. 

 CL CL2 RHOD RHOB RHOS RHOC RHOFCR AR ARS NS DFCR 

 1.1 1.1 0.300 0.300 0.160 0.300 0.000 0.80 1.20 112 3.00 

 

   WARNING FOLLOWING STAGE DESIGN LIMIT EXCEEDED ** *** 

 STAGE HUBTIP RATIO IS 0.95 DES LIMIT IS 0.94 

 ** HUB TIP RATIO IS TOO HIGH REDUCE HUB TIP RATIO INPUT ** 

 

   WARNING FOLLOWING STAGE DESIGN LIMIT EXCEEDED ** *** 

 STAGE BLADE HEIGHT IS 0.39 DES LIMIT IS 0.40 

 ** STAGE BLADE HEIGHT IS TOO SMALL CHANGE DES OPR OR REDUCE H/T INPUT ** 

 PR DEL H MACH AREA R HUB R TIP NB UTIPMAX STR WEIGHT TIN TMAX 

STAGE I 

 1.1448 14.7 0.312 0.131 7.45 7.84 84 1121.4 6292. 17. 1403. 1406. 

323. 

 COMPRESSOR COMPONENT WEIGHT SUMMARY 

 WTDSK WTBLD WTDRM WTST/IGV WTSF WTCAS WTNAB WTCR 

 109.2 10.4 10.9 17.5 9.4 55.7 9.5 0.0 

 N STG WEIGHT LENGTH CENGRA INERTIA LENGTH2 

 14 262.49 33.90 14.8 2797.8 33.90 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.310 566. 1460. 50401. 47255. 0.1182 1.3526 

 PR AD EF PO TO HP 

 15.0000 0.8600 50400.9 1459.6 12028. 

 HI HO WI CWI 

 152.54 358.53 41.27 28.82 

 ******************* TOTAL COMP WEIGHT IS 262.489 

 

 ************* 

 * * 

 * DUCT 6 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 DUCT , 6 

 RH= 7.53 RT= 7.83 LENG= 1.21 

 AREA= 0.102 RHO=.286 
NASA/TM—2005-213369 67 

 WTC(OUTER) WTC(INNER) WTT(REV) WT(TOTAL) TMIN 

 1.0137 0.9744 0.0000 1.9881 0.0593 

 

 ************* 

 * * 

 * PBUR 7 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 **************************************** 

 ALT MN VALUE 

 PTOT 0. 0.000 348.1 LB/SQIN 

 TTOT 0. 0.000 1465.0 DEG R 

 CWIN 0. 0.000 2.6 LB/SEC 

 **************************************** 

 BURNER NUMBER 7 

 RIN ROUT LENGTH MACH WSPEC TMIN 

 5.772 9.197 7.800 0.027 2.282 0.080 

 CAS WT INC WT LIN WT NOZ WT FRAME WTOT 

 10.5 6.6 16.2 4.1 76.8 129.8 

 HPC STRUCTURAL CASE: WEIGHT LENGTH DENSITY 

 15.6 12.1 0.160 

 

 ************* 

 * * 

 * HPT 8 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 **************************************** 

 ALT MN VALUE 

 PTOT 0. 0.000 332.5 LB/SQIN 

 TTOT 0. 0.000 2792.1 DEG R 

 CWOUT 39000. 0.800 16.4 LB/SEC 

 **************************************** 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.100 245. 2693. 47411. 47105. 0.4940 1.2957 

 UTIPMAX STRESS DEN W/AREA TR H/T 

 1353.5 23726.3 0.281 0.356 1.000 0.868 

 TURBINE 8 MECHANICAL DESIGN 

 H/T N STG LOADING AREA GE LOADING 

 0.868 2.000 0.340 0.494 0.843 

 UT RTIP RHUB DEL H RPM MAXRPM TORQ 

 1353.5 9.6 8.3 215.3 16202.9 16202.9 46992. 

 STAGE 1 

 WDISK WISD WID WBLADE WVANE WNB WCASE AR ARS TLBLADE TLVANE 

RHOB RHOD RHOS RHOC NS 

 34.1 11.9 0.0 2.5 4.7 2.4 10.7 1.20 1.20 1.06 1.06 

0.281 0.320 0.280 0.280 52 

 WSSFX WTRS WTSHRD WDRUMX WTRSHD WTDSB WTACC 

 2.7 1.1 2.3 6.2 0.0 0.0 11.7 
NASA/TM—2005-213369 68 

 PR DEL H MACH AREA R HUB R TIP NB MAXUTIP STR WEIGHT LENGTH 

STAGE I 

 2.0133 107.6 0.100 0.494 8.31 9.57 40 1353.5 23726. 90.24 2.83 

1348. 

 STAGE 2 

 WDISK WISD WID WBLADE WVANE WNB WCASE AR ARS TLBLADE TLVANE 

RHOB RHOD RHOS RHOC NS 

 22.4 7.8 0.0 1.5 2.8 1.5 6.8 1.80 1.80 0.67 0.67 

0.281 0.320 0.280 0.280 82 

 WSSFX WTRS WTSHRD WDRUMX WTRSHD WTDSB WTACC 

 1.7 0.7 1.5 3.9 0.0 0.0 7.4 

 PR DEL H MACH AREA R HUB R TIP NB MAXUTIP STR WEIGHT LENGTH 

STAGE I 

 2.2697 107.6 0.200 0.469 8.31 9.51 63 1345.1 22537. 57.99 1.57 

876. 

 TURBINE COMPONENT SUMMARY 

 WTDSK WTISD WTID WTDSB WTBLD WTRSH WTDRM WTRBS WTST 

WTSSH WTSF WTCAS WTNAB WTCC 

 56.4 19.7 0.0 0.0 4.0 0.0 10.1 1.9 7.5 

3.8 4.3 17.5 3.9 19.1 

 FRAME WT = 44.91 

 N STG LENGTH WEIGHT CENGRA INERTIA 

 2 4.40 193.13 3.7 2224. 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.300 624. 1944. 10355. 9764. 0.6897 1.3166 

 ***** AN**2 = 26.1 (BILLIONS IN2-RPM2) ***** 

 PR TR AD EF PO TO TO.1 

 4.5785 1.3739 0.9100 10355.2 1960.3 1943.5 

 H IN H OUT AREA FLOW HP 

 720.40 505.13 1.47 39.67 12081. 

 ******************* TOTAL TURB WEIGHT IS 193.134 

 

 ************* 

 * * 

 * DUCT 9 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 DUCT , 9 

 RH= 8.25 RT= 10.00 LENG= 1.39 

 AREA= 0.694 RHO=.286 

 WTC(OUTER) WTC(INNER) WTT(REV) WT(TOTAL) TMIN 
NASA/TM—2005-213369 69 

 1.2524 1.0340 0.0000 2.2864 0.0500 

 

 ************* 

 * * 

 * LPT 10 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 **************************************** 

 ALT MN VALUE 

 PTOT 0. 0.000 72.0 LB/SQIN 

 TTOT 0. 0.000 1943.6 DEG R 

 CWOUT 35000. 0.700 63.6 LB/SEC 

 **************************************** 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.165 345. 1944. 10306. 10124. 1.2154 1.3166 

 UTIPMAX STRESS DEN W/AREA TR H/T 

 808.2 15638.4 0.286 0.821 1.000 0.742 

 TURBINE 10 MECHANICAL DESIGN 

 H/T N STG LOADING AREA GE LOADING 

 0.742 3.000 0.280 1.215 1.176 

 UT RTIP RHUB DEL H RPM MAXRPM TORQ 

 808.2 11.1 8.3 139.8 8312.4 8312.4 61667. 

 STAGE 1 

 WDISK WISD WID WBLADE WVANE WNB WCASE AR ARS TLBLADE TLVANE 

RHOB RHOD RHOS RHOC NS 

 20.2 7.1 0.0 11.2 14.4 2.1 11.2 2.80 2.80 1.02 1.02 

0.286 0.320 0.280 0.280 75 

 WSSFX WTRS WTSHRD WDRUMX WTRSHD WTDSB WTACC 

 3.0 1.3 2.2 5.5 0.0 0.0 0.0 

 PR DEL H MACH AREA R HUB R TIP NB MAXUTIP STR WEIGHT LENGTH 

STAGE I 

 1.4961 46.6 0.165 1.215 8.27 11.14 83 808.2 15638. 78.39 2.54 

1491. 

 STAGE 2 

 WDISK WISD WID WBLADE WVANE WNB WCASE AR ARS TLBLADE TLVANE 

RHOB RHOD RHOS RHOC NS 

 19.4 6.8 0.0 11.6 15.1 2.0 10.7 3.25 3.25 0.96 0.96 

0.286 0.320 0.280 0.280 82 

 WSSFX WTRS WTSHRD WDRUMX WTRSHD WTDSB WTACC 

 2.9 1.2 2.1 5.2 0.0 0.0 0.0 

 PR DEL H MACH AREA R HUB R TIP NB MAXUTIP STR WEIGHT LENGTH 

STAGE I 

 1.5564 46.6 0.216 1.339 8.27 11.39 90 826.3 17223. 77.09 2.38 

1499. 

 STAGE 3 

 WDISK WISD WID WBLADE WVANE WNB WCASE AR ARS TLBLADE TLVANE 

RHOB RHOD RHOS RHOC NS 

 20.0 7.0 0.0 15.0 19.2 2.1 11.6 3.70 3.70 0.99 0.99 

0.286 0.320 0.280 0.280 83 

 WSSFX WTRS WTSHRD WDRUMX WTRSHD WTDSB WTACC 

 3.1 1.3 2.2 5.4 0.0 0.0 0.0 
NASA/TM—2005-213369 70 

 PR DEL H MACH AREA R HUB R TIP NB MAXUTIP STR WEIGHT LENGTH 

STAGE I 

 1.6348 46.6 0.268 1.621 8.27 11.95 92 866.6 20858. 86.89 2.23 

1776. 

 TURBINE COMPONENT SUMMARY 

 WTDSK WTISD WTID WTDSB WTBLD WTRSH WTDRM WTRBS WTST 

WTSSH WTSF WTCAS WTNAB WTCC 

 59.6 20.9 0.0 0.0 37.9 0.0 16.1 3.9 48.7 

6.5 9.0 33.6 6.2 0.0 

 FRAME COMPONENTS INFORMATION 

 FRAME WT LENGTH GAP #FRAMES ARF RHOF 

 4.93 3.8 1.9 6 1.20 0.120 

 CASE WT HUB WT UP SUPP LW SUPP RHOS THSUP WTTOWER TS DIAM 

 7.48 4.80 0.00 9.28 0.120 0.20 0.00 1.00 

 FRAME WT = 26.49 

 N STG LENGTH WEIGHT CENGRA INERTIA 

 3 12.92 268.87 5.7 4766. 

 DUCT 

 M NO VEL T TOT P TOT P STAT AREA GAM 

 0.319 576. 1441. 2705. 2528. 2.1345 1.3401 

 ***** AN**2 = 21.2 (BILLIONS IN2-RPM2) ***** 

 PR TR AD EF PO TO TO.1 

 3.8103 1.3491 0.9200 2704.9 1440.6 1440.6 

 H IN H OUT AREA FLOW HP 

 499.95 360.17 5.93 41.13 8133. 

 ******************* TOTAL TURB WEIGHT IS 268.869 

 

 ************* 

 * * 

 * FMIX 11 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 LENGTH= 14.77 WEIGHT = 46.81 

 

 ************* 

 * * 

 * DUCT 12 * 

 * * 

 ************2 

 
NASA/TM—2005-213369 71 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 

 ************* 

 * * 

 * NOZ 13 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 NOZZLE 13 

 TOTAL NOZZLE COMPONENT WEIGHT = 67.49 

 NOZZ WEIGHT= 53.73 TOT LENGTH= 18.155 TR WT= 1.50 

 THROAT AREA= 689.66 EXIT AREA= 689.66 

 MIXER EJECTOR LENGTH = 0.00 TRANS DUCT LENGTH = 0.00 

 NOZZLE WEIGHT BREAKDOWN 

 OUTER WALL = 53.73 PLUG = 12.26 NOZZ DENS = .2810 

 INNER WALL = 0.00 MIXER-EJECTOR= 0.00 PLUG DENS = .2810 

 CONV INNER WALLS = 0.00 NOZ AC LIN = 0.00 CASE DENS = .2810 

 DIV INNER WALLS = 0.00 PLG AC LIN = 0.00 M/E DENS = .2810 

 2-D SIDEWALLS = 0.00 AC SPLT PLATE = 0.00 ACOU DENS = .0900 

 VAR AREA C & A = 0.00 

 VAR AREA MISC = 0.00 

 2-D TRANS DUCT= 0.00 

 

 ************* 

 * * 

 * DUCT 14 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 FRAME COMPONENTS INFORMATION 

 FRAME WT LENGTH GAP #FRAMES ARF RHOF 

 3.20 3.7 18.3 3 1.80 0.120 

 CASE WT HUB WT UP SUPP LW SUPP RHOS THSUP WTTOWER TS DIAM 

 5.29 3.46 0.00 11.60 0.120 0.10 0.00 1.00 

 FRAME WT = 23.55 

 DUCT , 14 
NASA/TM—2005-213369 72 

 RH= 12.52 RT= 19.12 LENG= 56.10 

 AREA= 4.557 RHO=.168 

 WTC(OUTER) WTC(INNER) WTT(REV) WT(TOTAL) TMIN 

 56.6164 37.0721 0.0000 117.2388 0.0500 

 

 ************* 

 * * 

 * DUCT 15 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 

 ************* 

 * * 

 * SHAF 21 * 

 * * 

 ************2 

 

 MAX TORQUE CONDITION 

 *********************** 

 TORQUE 

 0.7 

 *********************** 

 SHAFT 21 

 DO DI LENG DN WT 

 2.24 2.07 13.41 0.92 6.6 

 BEARING # 3 4 

 BEARING WT 3.8 4.8 

 TOTAL SHAFT WEIGHT = 15.22 

 TOTAL INERTIA OF THIS SPOOL IS 5022. 

 

 ************* 

 * * 

 * SHAF 22 * 

 * * 

 ************2 

 

 MAX TORQUE CONDITION 

 *********************** 

 TORQUE 

 2.7 

 *********************** 

 SHAFT 22 

 DO DI LENG DN WT 

 1.67 1.00 64.31 0.35 49.1 

 BEARING # 1 2 5 

 BEARING WT 3.2 3.2 4.0 

 TOTAL SHAFT WEIGHT = 59.56 
NASA/TM—2005-213369 73 

 TOTAL INERTIA OF THIS SPOOL IS 11490. 

 

 ************* 

 * * 

 * INLT 1 * 

 * * 

 ************2 

 

 MAX CONDITIONS OCCUR AT 

 *********************** 

 ALT MN 

 PTOT 0. 0.000 

 TTOT 0. 0.000 

 *********************** 

 INLET 1 

 MDA SUBSONIC LOW DRAG INLET 

 INLET WEIGHT 

 WTINLT = 0.00 DUCTWT = 0.00 

 BDRWT = 0.00 TDRWT = 0.00 

 ENG MT = 58.55 FIRE WL= 0.00 

 FAN DIA= 38.46 LDUCT = 0.00 LDUCTS = 0.00 

 INLET LENGTH = 19.23 

 NACELLE WEIGHT 

 NAC WT = 191.18 

 INL CWL WT = 49.10 INLET LEN = 18.93 

 FAN CWL WT = 4.27 FAN BLD LEN = 2.96 

 FAN EXH CWL WT = 126.02 FAN EXH LEN = 106.00 

 CORE CWL WT = 0.00 CORE CWL LEN= 0.00 CWL AVG DIAM = 14.82 

 ANTI-ICE WT = 3.98 

 ACOUSTIC WT = 0.00 

 BULKHEAD WT = 7.81 

 TOTAL INLET/NACELLE WEIGHT = 249.74 

 

 ************* 

 * * 

 * ACCS WT * 

 * * 

 ************2 

 ACCS WT= 141.722 

1 

 COMP WT COMP ACCU UPSTREAM RADIUS DOWNSTREAM RADIUS 

 NO EST LEN LEN RI RO RI RO RI RO RI RO 

NSTAGE 

 1 250. 19.2 0. 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 

0 

 2 294. 5.9 6. 7.5 19.2 0.0 0.0 10.4 19.2 0.0 0.0 

1 

 3 0. 0.0 6. 10.4 19.2 0.0 0.0 10.4 12.2 12.2 19.2 

0 

 4 33. 8.5 14. 10.5 12.2 0.0 0.0 10.5 12.2 0.0 0.0 

0 
NASA/TM—2005-213369 74 

 5 262. 33.9 48. 3.6 7.8 0.0 0.0 7.5 7.8 0.0 0.0 

14 

 6 2. 1.2 49. 7.5 7.8 0.0 0.0 7.5 7.8 0.0 0.0 

0 

 7 130. 7.8 57. 5.8 9.2 0.0 0.0 5.8 9.2 0.0 0.0 

0 

 8 193. 4.4 62. 8.3 9.6 0.0 0.0 8.3 10.0 0.0 0.0 

2 

 9 2. 1.4 63. 8.3 10.0 0.0 0.0 8.3 10.0 0.0 0.0 

0 

 10 269. 12.9 76. 8.3 11.1 0.0 0.0 8.3 12.9 0.0 0.0 

3 

 11 47. 14.8 91. 8.1 13.0 13.0 18.2 8.1 18.2 0.0 0.0 

0 

 12 0. 0.0 91. 8.1 18.2 0.0 0.0 8.1 18.2 0.0 0.0 

0 

 13 67. 18.2 109. 8.3 18.2 0.0 0.0 0.0 14.8 0.0 0.0 

0 

 14 117. 56.1 76. 12.5 19.1 0.0 0.0 12.5 19.1 0.0 0.0 

0 

 15 0. 0.0 76. 8.3 12.9 0.0 0.0 8.3 12.9 0.0 0.0 

0 

 21 15. 13.4 0. 3.6 7.8 8.3 9.6 0.0 0.0 0.0 0.0 

0 

 22 60. 64.3 0. 7.5 19.2 8.3 11.1 0.0 0.0 0.0 0.0 

0 

 BARE ENGINE WEIGHT = 1491 

 ACCESSORIES WEIGHT = 141 

 TOTAL ENGINE WEIGHT = 1633 

 INLET/NACELLE WEIGHT = 249 

 TOTAL ENGINE POD WEIGHT = 1883 

 ENGINE LENGTH = 109.0 

 TOTAL ENGINE POD LENGTH = 128.2 

 ENGINE MAX DIAMETER = 38.5 

 NACELLE MAX DIAMETER = 45.8 

 ENGINE POD C.G. LOCATION = 33.4 

 ROTATING MACHINERY WEIGHT SUMMARY 

 COMP NAME TOT WT ROTOR STATOR FRAME LT LTM 

 2 FANH 294.4 112.6 125.6 56.2 19.9 14.9 

 5 HPC 262.5 140.1 82.6 39.7 33.9 28.6 

 8 HPT 193.1 94.2 54.1 44.9 4.4 4.4 

 10 LPT 268.9 140.7 101.7 26.5 12.9 7.1 

1 
NASA/TM—2005-213369 75 

Appendix C

ANSYS Axisymmetric Model Input File for Exoskeletal Rotor 

 This appendix contains the input data for an ANSYS structural analysis of the exoskeletal rotor when 

it is spinning at the overspeed condition of 16 400 rpm. The centrifugal force from each row of blades is 

represented by discrete force applied to the rotor. This model was used to generate the analysis results 

discussed in the section Drum Rotor Investigation. The ANSYS computer program is a large-scale 

multipurpose finite-element program that may be used for solving several classes of engineering 

analyses. The program contains many special features, one of which allows the two-dimensional 

modeling of a solid of revolution or an axisymmetric body. The use of axisymmetry reduces the size of 

the model by simplifying the generation of the geometry, structural loads, and boundary conditions. 

 This input file contains preprocessing, solution, and postprocessing commands to create the model, 

perform the analysis, and plot the results. The general preprocessor (PREP7) section contains solid 

modeling and mesh generation commands and is also used to define all other analysis data (geometric 

properties (real constants), material properties, etc.). Parameters are used in the geometry definition to 

provide more flexibility for the analyst to change the design. Loads and constraints are defined in the 

solution section (SOLUTION) where the analysis is also executed. The analysis results are reviewed 

using the postprocessor (POST1), in which plots are created to display distorted geometries, stress 

contours, and margins of safety contours. All plots of the geometry, mesh, and results are stored in a 

graphics file (file.grph), which may be viewed with the ANSYS Display utility. Another analyst may 

re-create the model by inputting this listing in either an ANSYS interactive session or a batch job. 

 The following ANSYS FEA input listing was generated by Daniel N. Kosareo. 

/BATCH 

! 

! ANSYS RELEASE= 7.0 UP20021010 01/08/03 Daniel N. Kosareo 

! 

/SHOW,file,grph 

! 

! DEFINE PARAMETERS 

! 

*ASK,DTHK,BLADE MOUNTING RING THICKNESS,0.375 

*ASK,RTHK,ROTOR THICKNESS,0.125 

*ASK,FHGT,FLANGE HEIGHT,0.5 

! 

! PREPROCESSING MODULE 

! 

/PREP7 

/TITLE,Exoskeletal AE3007 Engine, Rotor Design, 16400 RPM 

/WIN,1,-1,1,-.88,1 

/VIEW,1,,,1 

/VUP,1,X 

! 

! DEFINE ELEMENT TYPES (Plane Axisymmetric Elements) 
NASA/TM—2005-213369 76 

! 

ET,1,PLANE82,,,1 

*REPEAT,21,1 

! 

! MATERIAL PROPERTIES FOR GRAPHITE-POLYIMIDE COMPOSITE Gr/P 

! 

MP,EX,1,24.0E+06 

MP,NUXY,1,0.33 

MP,DENS,1,(0.056/386.4) 

MP,ALPX,1,20E-06 

! 

! MATERIAL PROPERTIES FOR TITANIUM Ti-6Al-4V 

! 

MP,EX,2,16.0E+06 

MP,NUXY,2,0.31 

MP,DENS,2,(0.160/386.4) 

MP,ALPX,2,4.9E-06 

! 

! MATERIAL PROPERTIES FOR HASTALLOY X 

! 

MP,EX,3,29.8E+06 

MP,NUXY,3,0.32 

MP,DENS,3,(0.297/386.4) 

MP,ALPX,3,7.5E-06 

! 

! MATERIAL PROPERTIES FOR SILICON-CARBIDE SiC/SiC 

! 

MP,EX,4,10.9E+06 

MP,NUXY,4,0.07 

MP,DENS,4,(0.08/386.4) 

MP,ALPX,4,1.7E-06 

! 

! SOLID AXISYMMETRIC MODEL 

! 

K,1,7.84,0.00 

K,2,7.84,2.25 ! Stage 1 

K,4,7.84,3.59 

KFILL,2,4 

K,5,7.84,4.72 ! Stage 2 

K,7,7.84,5.79 

KFILL,5,7 

K,8,7.84,7.24 ! Stage 3 

K,10,7.84,8.15 

KFILL,8,10 

K,11,7.84,9.276 ! Stage 4 

K,13,7.84,10.19 

KFILL,11,13 

K,14,7.84,11.314 ! Stage 5 

K,16,7.84,12.01 

KFILL,14,16 
NASA/TM—2005-213369 77 

K,17,7.84,13.19 ! Stage 6 

K,19,7.84,13.78 

KFILL,17,19 

K,20,7.84,14.585 ! Stage 7 

K,22,7.84,15.23 

KFILL,20,22 

K,23,7.84,16.02 ! Stage 8 

K,25,7.84,16.41 

KFILL,23,25 

K,26,7.84,17.37 ! Stage 9 

K,28,7.84,17.80 

KFILL,26,28 

K,29,7.84,18.60 ! Stage 10 

K,31,7.84,19.09 

KFILL,29,31 

K,32,7.84,19.86 ! Stage 11 

K,34,7.84,20.32 

KFILL,32,34 

K,35,7.84,21.1 ! Stage 12 

K,37,7.84,21.66 

KFILL,35,37 

K,38,7.84,22.38 ! Stage 13 

K,40,7.84,22.74 

KFILL,38,40 

K,41,7.84,23.55 ! Stage 14 

K,43,7.84,23.92 

KFILL,41,43 

K,44,7.84,24.68 

K,45,7.84+0.5*(10.29-7.84)/(32.4-24.68),24.68+0.5 

K,46,7.84+(32.4-24.68-0.5)*(10.29-7.84)/(32.4-24.68),32.4-0.5 

K,47,10.29,32.4 

K,48,10.29+0.5*(9.57-10.29)/(39.04-32.4),32.4+0.5 

K,49,10.29+(39.04-32.4-0.5)*(9.57-10.29)/(39.04-32.4),39.04-0.5 

K,50,9.57,39.04 ! HPT Stage 1 

K,52,9.57,39.95 

KFILL,50,52 

K,53,9.51,41.4 ! HPT Stage 2 

K,55,9.51,42.25 

KFILL,53,55 

K,56,9.51,42.25+0.5 

L,1,2 

*REPEAT,55,1,1 

K,57,7.84+DTHK,0.00 

K,58,7.84+DTHK,3.59 

K,59,7.84+DTHK,5.79 

K,60,7.84+DTHK,8.15 

K,61,7.84+DTHK,10.19 

K,62,7.84+DTHK,12/01 

K,63,7.84+DTHK,13.78 

K,64,7.84+DTHK,15.23 
NASA/TM—2005-213369 78 

K,65,7.84+DTHK,16.41 

K,66,7.84+DTHK,17.80 

K,67,7.84+DTHK,19.09 

K,68,7.84+DTHK,20.32 

K,69,7.84+DTHK,21.66 

K,70,7.84+DTHK,22.74 

K,71,7.84+DTHK,23.92 

KGEN,2,44,49,1,DTHK,,, 

K,78,9.57+DTHK,39.04 

K,79,9.57+DTHK,39.95 

K,80,9.51+DTHK,42.25 

K,81,9.51+DTHK,42.25+0.5 

L,57,58 

*REPEAT,24,1,1 

L,1,57 

*REPEAT,15,3,1 

L,44,72 

*REPEAT,7,1,1 

L,52,79 

L,55,80 

L,56,81 

KGEN,2,57,63,1,RTHK,,, 

K,89,KX(64)+RTHK,15.23-0.5 

K,90,KX(64)+RTHK,15.23 

K,91,KX(64)+RTHK,15.23+0.5 

KGEN,2,65,71,1,RTHK,,, 

K,99,KX(71)+RTHK,24.68-0.5 

KGEN,2,72,78,1,RTHK,,, 

K,107,KX(78)+RTHK,39.04+0.5 

KGEN,2,79,81,1,RTHK,,, 

L,82,83 

*REPEAT,28,1,1 

L,57,82 

*REPEAT,7,1,1 

L,64,90 

L,65,92 

*REPEAT,7,1,1 

L,72,100 

*REPEAT,7,1,1 

L,79,108 

L,80,109 

L,81,110 

KGEN,2,89,91,1,FHGT,,, 

KGEN,2,99,101,1,FHGT,,, 

KMODIF,114,KX(116) 

KMODIF,115,KX(116) 

KGEN,2,105,107,1,FHGT,,, 

KMODIF,118,KX(117) 

KMODIF,119,KX(117) 

L,111,112 
NASA/TM—2005-213369 79 

L,112,113 

L,89,111 

*REPEAT,3,1,1 

L,114,115 

L,115,116 

L,99,114 

*REPEAT,3,1,1 

L,117,118 

L,118,119 

L,105,117 

*REPEAT,3,1,1 

/PNUM,KP,1 

KPLO 

LPLO 

/PNUM,KP,0 

/NUM,-1 

/DEVICE,VECT,ON 

LPLO 

/NUM,1 

/DEVICE,VECT,OFF 

/DEVICE,RAST,ON 

LPLO 

/COM 

KSEL,S,KP,,1,4,1 

KSEL,A,KP,,57,58 

LSLK,S,1 

AL,ALL ! Area 1 

KSEL,S,KP,,4,7,1 

KSEL,A,KP,,58,59 

LSLK,S,1 

AL,ALL ! Area 2 

KSEL,S,KP,,7,10,1 

KSEL,A,KP,,59,60 

LSLK,S,1 

AL,ALL ! Area 3 

KSEL,S,KP,,10,13,1 

KSEL,A,KP,,60,61 

LSLK,S,1 

AL,ALL ! Area 4 

KSEL,S,KP,,13,16,1 

KSEL,A,KP,,61,62 

LSLK,S,1 

AL,ALL ! Area 5 

KSEL,S,KP,,16,19,1 

KSEL,A,KP,,62,63 

LSLK,S,1 

AL,ALL ! Area 6 

KSEL,S,KP,,19,22,1 

KSEL,A,KP,,63,64 

LSLK,S,1 
NASA/TM—2005-213369 80 

AL,ALL ! Area 7 

KSEL,S,KP,,22,25,1 

KSEL,A,KP,,64,65 

LSLK,S,1 

AL,ALL ! Area 8 

KSEL,S,KP,,25,28,1 

KSEL,A,KP,,65,66 

LSLK,S,1 

AL,ALL ! Area 9 

KSEL,S,KP,,28,31,1 

KSEL,A,KP,,66,67 

LSLK,S,1 

AL,ALL ! Area 10 

KSEL,S,KP,,31,34,1 

KSEL,A,KP,,67,68 

LSLK,S,1 

AL,ALL ! Area 11 

KSEL,S,KP,,34,37,1 

KSEL,A,KP,,68,69 

LSLK,S,1 

AL,ALL ! Area 12 

KSEL,S,KP,,37,40,1 

KSEL,A,KP,,69,70 

LSLK,S,1 

AL,ALL ! Area 13 

KSEL,S,KP,,40,43,1 

KSEL,A,KP,,70,71 

LSLK,S,1 

AL,ALL ! Area 14 

ALLSEL 

A,43,44,72,71 ! Area 15 

*REPEAT,7,1,1,1,1 ! Areas 16-21 

KSEL,S,KP,,50,52,1 

KSEL,A,KP,,78,79 

LSLK,S,1 

AL,ALL ! Area 22 

KSEL,S,KP,,52,55,1 

KSEL,A,KP,,79,80,1 

LSLK,S,1 

AL,ALL ! Area 23 

ALLSEL 

A,55,56,81,80 ! Area 24 

A,57,58,83,82 ! Area 25 

*REPEAT,6,1,1,1,1 ! Areas 26-30 

KSEL,S,KP,,88,90,1 

KSEL,A,KP,,63,64 

LSLK,S,1 

AL,ALL ! Area 31 

KSEL,S,KP,,90,92,1 

KSEL,A,KP,,64,65 
NASA/TM—2005-213369 81 

LSLK,S,1 

AL,ALL ! Area 32 

ALLSEL 

A,65,66,93,92 ! Area 33 

*REPEAT,6,1,1,1,1 ! Areas 34-38 

KSEL,S,KP,,98,100,1 

KSEL,A,KP,,71,72 

LSLK,S,1 

AL,ALL ! Area 39 

ALLSEL 

A,72,73,101,100 ! AREA 40 

*REPEAT,6,1,1,1,1 ! Areas 41-45 

KSEL,S,KP,,106,108 

KSEL,A,KP,,78,79 

LSLK,S,1 

AL,ALL ! AREA 46 

ALLSEL 

A,79,80,109,108 ! AREA 47 

A,80,81,110,109 ! AREA 48 

A,89,90,112,111 ! Area 49 

A,90,91,113,112 ! Area 50 

A,99,100,115,114 ! Area 51 

A,100,101,116,115 ! Area 52 

A,105,106,118,117 ! Area 53 

A,106,107,119,118 ! Area 54 

ADEL,16,21,1,1 

ASEL,S,AREA,,25,30,1 

AADD,ALL ! AREA 16 

ASEL,S,AREA,,31,49,18 

AADD,ALL ! AREA 17 

ASEL,S,AREA,,32,39,1 

ASEL,A,AREA,,50,51,1 

ASEL,A,AREA,,15,15,1 

AADD,ALL ! AREA 18 

ASEL,S,AREA,,40,45,1 

ASEL,A,AREA,,52,53,1 

AADD,ALL ! AREA 15 

ASEL,S,AREA,,46,48,1 

ASEL,A,AREA,,54,54,1 

ASEL,A,AREA,,24,24,1 

AADD,ALL ! AREA 19 

ALLSEL 

NUMCMP,AREA 

/PNUM,AREA,1 

APLO 

/PNUM,AREA,0 

/PNUM,KP,1 

APLO 

/PNUM,KP,0 

/PNUM,AREA,1 
NASA/TM—2005-213369 82 

APLO 

APLO,1,14,1 

APLO,15,19,1 

APLO,20,21,1 

/PNUM,AREA,0 

*DO,ANUM,1,7,1 

ASEL,S,AREA,,ANUM,ANUM,1 

AATT,1,1,ANUM 

*ENDDO 

*DO,ANUM,8,14,1 

ASEL,S,AREA,,ANUM,ANUM,1 

AATT,2,1,ANUM 

*ENDDO 

ASEL,S,AREA,,15 

AATT,3,1,15 

ASEL,S,AREA,,16 

AATT,1,1,16 

ASEL,S,AREA,,17 

AATT,2,1,17 

ASEL,s,AREA,,18 

AATT,2,1,18 

ASEL,S,AREA,,19 

AATT,3,1,19 

ASEL,S,AREA,,20 

AATT,4,1,20 

ASEL,S,AREA,,21 

AATT,4,1,21 

ALLSEL 

/pnum,mat,1 

aplo 

/pnum,mat,1 

! 

! AUTOMATED FINITE ELEMENT MESHING 

! 

esize,RTHK/2 

MSHAPE,0,2D 

MSHKEY,2 

amesh,all 

/pnum,type,1 

eplo 

/pnum,type,0 

KSEL,S,KP,,1,57,56 

KSEL,A,KP,,82,82,1 

LSLK,S,1 

NSLL,S,1 

CP,1,UY,ALL 

KSEL,S,KP,,56,81,81-56 

KSEL,A,KP,,110,110,1 

LSLK,S,1 

NSLL,S,1 
NASA/TM—2005-213369 83 

CP,2,UY,ALL 

ALLSEL 

FINI 

! 

! SOLUTION MODULE 

! 

/SOLU 

ANTYPE,STATIC,NEW 

OUTRES,ALL,ALL 

TIME,1.0 

! 

! ROTATIONAL SPEED 

! 

ROTSPD=2*PI*16400/60 

OMEGA,0,ROTSPD,0,0 

! 

! DISPLACEMENT CONSTRAINTS 

! 

DK,1,UY,0 

! DK,82,UX,0 

! DK,110,UX,0 

! 

! CENTRIFUGAL FORCES DUE TO WEIGHT OF BLADES AT EACH ROW 

! 

FK,3,FX,21*0.05825*7.84*ROTSPD*ROTSPD/386.4 

FK,6,FX,28*0.02382*7.84*ROTSPD*ROTSPD/386.4 

FK,9,FX,35*0.01176*7.84*ROTSPD*ROTSPD/386.4 

FK,12,FX,42*0.0064084*7.84*ROTSPD*ROTSPD/386.4 

FK,15,FX,49*0.0038977*7.84*ROTSPD*ROTSPD/386.4 

FK,18,FX,56*0.0024879*7.84*ROTSPD*ROTSPD/386.4 

FK,21,FX,62*0.0016691*7.84*ROTSPD*ROTSPD/386.4 

FK,24,FX,68*0.0033844*7.84*ROTSPD*ROTSPD/386.4 

FK,27,FX,74*0.0025427*7.84*ROTSPD*ROTSPD/386.4 

FK,30,FX,78*0.0019597*7.84*ROTSPD*ROTSPD/386.4 

FK,33,FX,82*0.0015143*7.84*ROTSPD*ROTSPD/386.4 

FK,36,FX,84*0.0013144*7.84*ROTSPD*ROTSPD/386.4 

FK,39,FX,85*0.0011129*7.84*ROTSPD*ROTSPD/386.4 

FK,42,FX,84*0.0010030*7.84*ROTSPD*ROTSPD/386.4 

/PBC,ALL,,1 

APLO 

/PBC,ALL,,0 

SBCTRA 

/PBC,ALL,,1 

EPLO 

/PBC,ALL,,0 

LSWRITE,1 

LSSOLVE,1 

FINI 

! 

! POSTPROCESSING 
NASA/TM—2005-213369 84 

! 

/POST1 

SET,1,1 

/DSCALE,1,1.0 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of Radial Displacements (in) 

PLNS,U,X 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of Radial Stresses (Sx) 

PLNS,S,X 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of Axial Stresses (Sy) 

PLNS,S,Y 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of Tangential (Hoop) Stresses (Sz) 

PLNS,S,Z 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of 1st Principal Stress (S1) 

PLNS,S,1 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of 2nd Principal Stress (S2) 

PLNS,S,2 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of 3rd Principal Stress (S3) 

PLNS,S,3 

/ANNOT,DELE 

/ANNOT,ON 

/TSPEC,15,.9,2,0. 

/TLABEL,-.9,.94,Contour Plot of Max. Equivalent Stress (Seqv) 

PLNS,S,EQV 

/ANNOT,DELE 

FINI 

/EXIT 
NASA/TM—2005-213369 85 

Appendix D 

Blade Profiles for Exoskeletal Rotor 

 This appendix contains the input data for the Pro/Engineer® models of the compressor and turbine 

rotor blades. The exoskeletal compressor rotor blade shapes are based on the NACA 65A010 compressor 

airfoil shape. The NACA 65A010 profile is a symmetric shape specifically developed for compressor 

applications. This profile is shown in figure 50 and is listed in table 22. This airfoil has a 10-percentthickness-to-chord ratio. 

TABLE 22.⎯NACA 65A010 COMPRESSOR AIRFOIL 

[Leading-edge radius, 0.00636c; trailing-edge radius, 0.00023c.] 

Length-tochord ratio, 

x/c, 

percent 

Thickness-tochord ratio, 

y/c, 

percent 

Length-tochord ratio, 

x/c, 

percent 

Thickness-tochord ratio, 

y/c, 

percent 

0 0.0 40 4.995 

.5 .765 45 4.983 

.75 .928 50 4.863 

1.25 1.183 55 4.632 

2.5 1.623 60 4.304 

5 2.182 65 3.809 

7.5 2.65 70 3.432 

10 3.04 75 2.912 

15 3.658 80 2.352 

20 4.127 85 1.771 

25 4.483 90 1.188 

30 4.742 95 .604 

35 4.912 100 0 

 A stagger angle θ is used for the base and the tip of the compressor blades, which gives the blade a 

twist along its length. The stagger angle is applied at the centroid of the airfoil, as shown in figure 51. 

In the high-pressure turbine, the airfoil camber line is curved, and the symmetric shape of the NACA 

65A010 profile is distributed about the camber line. The curved camber line changes the fluid flow to the 

desired direction. In the case of the exoskeletal design, the curved camber line is a circular arc as shown 

in figure 52. The stagger angle θ of the curved airfoil is the average of the camber line inlet angle φi and 

exit angle φe. From basic trigonometry, the radius r of the camber line is given by the following equation: 
NASA/TM—2005-213369 86 

sin

cos cos e i

c

r

θ = φ− φ (D1) 

where c is the chord. 

 In the Pro/Engineer® model, each point of the NACA 65A010 profile has a local radius rn and a local 

camber line angle φn, which are defined by the following set of equations: 

2 n

x

c

⎛ ⎞ ∆φ = θ⎜ ⎟ ⎝ ⎠ (D2a) 

n

y r rc

c

⎛ ⎞ = + ⎜ ⎟ ⎝ ⎠ (D2b) 

90 n in φ = + φ − ∆φ (D2c) 

The coordinates for each point on the profile are given by the following transformations: 

cos

2 nn n

c

x r ′ = φ+ (D3) 

 

2 sin

tan nn n

c

y r ′ = φ− θ

 (D4) 
NASA/TM—2005-213369 87 

References 

 1. Mattingly, Jack D.: Elements of Gas Turbine Propulsion. McGraw-Hill, New York, NY, 1996. 

 2. Halliwell, Ian: Exoskeletal Engine Concept: Feasibility Studies for Medium and Small Thrust 

Engines. NASA/CR⎯2001-211322, 2002. http://gltrs.grc.nasa.gov/reports/2001/CR–2001–

211322.pdf 

 3. Gunston, Bill, ed.: Jane’s Aero-Engines. Issue 14, Jane’s Information Group Limited, Coulsdon, 

Surrey, 2003. 

 4. Tong, Michael T., et al.: A Computer Code for Gas Turbine Engine Weight and Disk Life Estimation. 

ASME GT−2002−30500, 2002. 

 5. Onat, E.; and Klees, G.W.: A Method to Estimate Weight and Dimensions of Large and Small Gas 

Turbine Engines; Final Report. NASA CR−159481, 1979. 

 6. Pro/ENGINEER®, PTC Worldwide Headquarters, Needham, MA, 2001. 

 7. Maintaining U.S. Leadership in Aeronautics: Breakthrough Technologies to Meet Future Air and 

Space Transportation Needs and Goals. Contract No. NASW−4938, National Academy Press, 

Washington, DC, 1998. 

 8. Klann, John L.; and Snyder, Christopher A.: NEPP Programmers Manual (NASA Engine 

Performance Program). NASA TM−106575−VOL−1, 1994. Available from the NASA Center for 

Aerospace Information. 

 9. Converse, G.L.; and Giffin, R.G.: Extended Parametric Representation of Compressor Fans and 

Turbines. Volume 1: CMGEN User’s Manual; Final Report. NASA CR−174645, 1984. 

10. Converse, G.L.: Extended Parametric Representation of Compressor Fans and Turbines. Volume 2: 

Part User’s Manual (Parametric Turbine); Final Report. NASA CR−174646, 1984. 

11. Onat, E.; and Klees, G.W.: A Method to Estimate Weight and Dimensions of Large and Small Gas 

Turbine Engines; Final Report. NASA CR−159481, 1979. 

12. Tong, Michael, et al.: A Computer Code for Gas Turbine Engine Weight and Disk Life Estimation. 

ASME GT−2002−30500, 2002. 

13. Abumeri, Galib H.; Kuguoglu, Latife H.; and Chamis, Christos C.: Composite Fan Blade Design for 

Advanced Engine Concepts. NASA/TM⎯2004-212943, 2004. 

14. Kuguoglu, Latife; Abumeri, Galib; and Chamis, Christos C.: Structural Evaluation of Exo-Skeletal 

Engine Fan Blades. AIAA 2003−1861, 2003. 

15. MSC.Nastran 2001. MSC Software Corporation, Los Angeles, CA, 2001. 

16. Metallic Materials and Elements for Aerospace Vehicle Structures. MIL−HDBK−5F, 1990. 

17. HyperSizerTM. Structural Sizing Software, Collier Research Corporation, Hampton, VA. 

http://www.collier-research.com/Products/products.htm Accessed Nov. 5, 2004. 

18. MSC.Patran 2002 (r2). Release Guide, MSC Software Corporation, Santa Ana, CA, 2002. 

19. ANSYS Verification Manual. Verification Test Case Descriptions, ©2002 SAS IP, Inc., 2002. 

20. Shigley, Joseph Edward: Mechanical Engineering Design. Second ed., McGraw-Hill, New York, NY, 

1972. 

21. ANSYS Technical Overview (MBA−A15−9/97), ANSYS, Inc., ©1997. 

22. Gunter, E.J.: Critical Speeds of a Flexible Rotor, Department of Mechanical Engineering, University 

of Virginia, Nov. 10, 1971.

23. Vance, John M.: Rotordynamics of Turbomachinery. Wiley, New York, NY, 1988. 

24. Timoshenko, S.; Young, D.H.; and Weaver, W., Jr.: Vibration Problems in Engineering. Wiley, New 

York, NY, 1974. 

25. Childs, Dara W.: Turbomachinery Rotordynamics: Phenomena, Modeling, and Analysis. Wiley, New 

York, NY, 1993. 

26. Prohl, M.A.: A General Method for Calculating Critical Speeds of Flexible Rotors. J. Appl. Mech. 

Trans. ASME, vol. 12, no. 3, 1945, pp. A142−A148. 
NASA/TM—2005-213369 88 

27. Myklestad, N.O.: A New Method of Calculating Natural Modes of Uncoupled Bending Vibration of 

Airplane Wings and Other Type of Beams. J. Aero. Sci., vol. 11, 1944, pp. 153−162. 

28. Linn, F.C.; and Prohl, M.A.: The Effect of Flexibility of Support Upon the Critical Speeds of HighSpeed Rotors. Trans. Soc. Naval Architects Mar. Eng., vol. 59, 1951, pp. 536−553. 

29. Ehrich, Frederic F.: Handbook of Rotordynamics. McGraw-Hill, New York, NY, 1992. 

30. DellaCorte, Chris; and Pinkus, Oscar: Tribological Limitations in Gas Turbine Engines: A Workshop 

to Identify the Challenges and Set Future Directions; Revised. NASA/TM⎯2000-210059/REV1, 

2002. http://gltrs.grc.nasa.gov/reports/2002/TM−2000−210059−REV1.pdf 

31. Sullivan, Roy: Thermal Analysis of a Carbon-Carbon Bearing Design for Exoskeletal Engine 

Bearings. NASA/TM⎯2001-210946, 2001. Available from the NASA Center for AeroSpace 

Information. 

32. Dellacorte, Christopher; and Valco, Mark J.: Load Capacity Estimation of Foil Air Journal Bearings 

of Oil-Free Turbomachinery Applications. NASA/TM—2000-209782, 2000. 

http://gltrs.grc.nasa.gov/reports/2000/TM-2000-209782.pdf
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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Technical Memorandum

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National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135–3191

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August 2005

NASA TM—2005-213369

E–14837

WBS–22–708–87–07

93

Investigation of Exoskeletal Engine Propulsion System Concept

Joseph M. Roche, Donald T. Palac, James E. Hunter, David E. Myers,

Christopher A. Snyder, Daniel N. Kosareo, David R. McCurdy,

and Kevin T. Dougherty

Propulsion system configurations; Turbofan engines; Jet engines; Turbine engines;

Gas turbine engines

Unclassified -Unlimited

Subject Category: 07 Distribution: Nonstandard

Joseph M. Roche, Donald T. Palac, James E. Hunter, David E. Myers, and Christopher A. Snyder, NASA Glenn

Research Center; Daniel N. Kosareo, ZIN Technologies, Inc., Brook Park, Ohio 44142; and David R. McCurdy and

Kevin T. Dougherty, QSS Group, Inc., Cleveland, Ohio 44135. Responsible person, Kevin T. Dougherty, organization

code PR, 216–433–3546.

An innovative approach to gas turbine design involves mounting compressor and turbine blades to an outer rotating

shell. Designated the exoskeletal engine, compression (preferable to tension for high-temperature ceramic materials,

generally) becomes the dominant blade force. Exoskeletal engine feasibility lies in the structural and mechanical design

(as opposed to cycle or aerothermodynamic design), so this study focused on the development and assessment of a

structural-mechanical exoskeletal concept using the Rolls-Royce AE3007 regional airliner all-axial turbofan as a

baseline. The effort was further limited to the definition of an exoskeletal high-pressure spool concept, where the major

structural and thermal challenges are represented. The mass of the high-pressure spool was calculated and compared with

the mass of AE3007 engine components. It was found that the exoskeletal engine rotating components can be significantly

lighter than the rotating components of a conventional engine. However, bearing technology development is required,

since the mass of existing bearing systems would exceed rotating machinery mass savings. It is recommended that once

bearing technology is sufficiently advanced, a “clean sheet” preliminary design of an exoskeletal system be accomplished

to better quantify the potential for the exoskeletal concept to deliver benefits in mass, structural efficiency, and cycle

design flexibility.