Gregory P. Herrick

Glenn Research Center, Cleveland, Ohio

Effects of Inlet Distortion on Aeromechanical 

Stability of a Forward-Swept High-Speed Fan

NASA/TM—2011-216959

January 2011

AIAA–2010–6711
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Gregory P. Herrick

Glenn Research Center, Cleveland, Ohio

Effects of Inlet Distortion on Aeromechanical 

Stability of a Forward-Swept High-Speed Fan

NASA/TM—2011-216959

January 2011

AIAA–2010–6711

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

46th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASEE

Nashville, Tennessee, July 25–28, 2010
Acknowledgments

The author gratefully acknowledges the funding support of the Turbomachinery portion of the Subsonic Fixed Wing Program. 

Nateri Madavan is the project manager, and Ruben DelRosario is the Program Associate Principal Investigator.

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This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-216959 1

Effects of Inlet Distortion on Aeromechanical Stability of a 

Forward-Swept High-Speed Fan

Gregory P. Herrick

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein 

the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present 

serious concerns with regard to aerodynamic and aeromechanic performance of the compression system 

in response to inlet distortion. Separately, a forward-swept high-speed fan was developed to address noise 

concerns of modern podded turbofans; however this fan encounters aeroelastic instability (flutter) as it

approaches stall. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is 

applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in 

its application to assess aerodynamic damping of vibrating blades at various flow conditions, is modified 

and then applied in a computational study to preliminarily assess the effects of inlet distortion on 

aeroelastic stability of the fan. Computational engineering application and implementation issues are 

discussed, followed by an investigation into the aeroelastic behavior of the fan with clean and distorted 

inlets.

Nomenclature

BLI Boundary-Layer-Ingesting

ECMF Exit Corrected Mass Flow (i.e., “exit throttle”) boundary condition

ND Nodal Diameter

#NDB (#) Nodal Diameter, Backward-traveling wave

#NDF (#) Nodal Diameter, Forward-traveling wave

NS Near-Stall

PE Peak-Efficiency

PRE Exit Static Pressure with Radial Equilibrium boundary condition

SPP Exit Static Pressure Profile boundary condition.

Introduction

Since the advent of turbo/jet propulsion, “clean” intake flow for the fan (or first stage) of the 

turbomachine propulsor has been the desired, elusive, goal of airframe, inlet, and propulsion system 

designers alike. Engines mounted at the aft extent of the fuselage ingest flow influenced by the airframe 

boundary layer. Engines mounted within the tail/aft-fuselage assembly ingest airframe boundary layer and 

then suffer additional flow degradation due to the serpentine inlet duct. Tail- and aft-fuselage-mounted 

engines were particularly common in 1960s commercial aircraft design, when propulsive efficiency, 

noise, and emissions were of lower precedence than today. While modern small subsonic transport

designs employ podded engines mounted on the aft-fuselage, these designs do not incorporate engines 

embedded within the tail assembly. Nearly all large subsonic transport designs since the 1980s have 

employed podded engines, mounted by pylon, on the wings. With this placement, the engines receive 

cleaner inlet flow than the aforementioned fuselage-mounted engines, facilitating more powerful, more 

cost-efficient engine designs than their predecessors.
NASA/TM—2011-216959 2

As we look to the future, design goals now focus on reduced emissions, reduced fuel consumption, 

better propulsive efficiency, and reduced ambient noise. Among the prominent design concepts under 

development for application to future aircraft is the “hybrid wing/body.” Most of these prospective 

designs feature propulsion systems wherein the propulsive fan, whether coupled directly to a Brayton 

cycle turbine engine or part of a turbo-electric distributed propulsion system, is embedded partially or 

fully submerged within the body.

The design process has been revolutionized through the aforementioned evolution of airframe, inlet, 

and propulsion system design. Analytical theory (often with simplifying assumptions) and repeating 

(costly) cycles of design, build, test have been supplemented greatly with advances in computational 

technology, both hardware and software. In this research effort, a parallel computational fluid dynamics 

(CFD) code is modified and customized to begin studying the issues relevant to aeromechanical response 

of turbomachinery to distorted inlet flows.

Background

A primary design concept of research interest in the United States is the hybrid wing/body. As its 

name implies, this design is a departure from the tube-and-wing designs which have populated the skies 

for the last 60 years. Liebeck (Ref. 1) published an authoritative paper on this airframe concept in 2002. 

This design melds the wing and body in an integral unit. Among the goals mentioned previously, greater 

propulsive efficiency is attained with this design by utilizing boundary layer ingestion (BLI). Studies have 

shown that wake-ingestion can increase propulsive efficiency by 20 percent (Smith (Ref. 2) and Plas et al.

(Ref. 3)). While BLI improves propulsion system efficiency, it can wreak havoc on turbomachinery

aerodynamics and aeromechanics, and hence high cycle fatigue (HCF), when unchecked. Kenyon et al.

(Ref. 4) described the significant role of total pressure flow distortion in the onset of non-uniform flowrelated HCF.

Computational research into inlet distortion concerns ramped up in the mid-1990s. Hah et al. (Ref. 5)

were pioneers in numerical analysis of inlet distortion, studying a 1/8th sector of a rotor with a total 

pressure distortion at the inlet plane. They prescribed constant static pressure at one location on the 

shroud at the exit rotor plane. Hirai et al. (Ref. 6) used full-annulus numerical simulations of a rotor 

subject to circumferentially non-uniform total pressure at the inlet boundary to investigate the mechanism 

of pressure loss associated with passage shock motion; they held their exit static pressure constant.

Charalambous et al. (Ref. 7) used CFD to study the changes in axial compressor performance due to 

circumferential and stratified distortions in inlet total pressure; they maintained a constant static pressure 

at the exit. Yao et al. (Ref. 8) and Gorell et al. (Ref. 9) used CFD to demonstrate that a 1/rev total pressure 

distortion at the inlet incites a static pressure distortion which in turn induces a swirl distortion; they 

enforced an interpolated, non-uniform pressure distribution about the exit plane. Bréard et al. (Ref. 10) 

computed forced response of a fan rotor subjected to inlet distortion; they specified mass flows at blade 

row interfaces. Zemp et al. (Ref. 11) conducted full-annulus, unsteady CFD simulations of a centrifugal 

compressor subject to inlet distortion; they applied a uniform average static pressure over the entire exit 

area which was iteratively adjusted to match computed mass flows with measured mass flows at the 

respective operating point.

Analysis Code

The relevant CFD code modified and applied in this study is TURBO. TURBO is a physics-based 

simulation tool for multistage turbomachinery. The solver computes the fluid conservation laws without 

ad hoc modeling of any flow phenomena other than models required for turbulence. This code solves the 

unsteady Reynolds-averaged Navier-Stokes equations and a decoupled k-ε turbulence model developed 

by Zhu and Shih (Ref. 12). To facilitate rotor-stator interaction, TURBO employs a sliding interface

technique implemented by Chen and Barter (Ref. 13) in which conservative variables are interpolated 
NASA/TM—2011-216959 3

across blade row interfaces. The code is implemented in a portable, scalable form for distributed-memory 

parallel computers using MPI message passing. The parallel implementation employs domain 

decomposition and supports general multi-block grids with arbitrary grid-block connectivity. The solution 

algorithm is a Newton iterative implicit time-accurate scheme with characteristics-based finite-volume 

spatial discretization. The Newton subiterations are solved using a concurrent block-Jacobi symmetric 

Gauss-Seidel (BJ-SGS) relaxation scheme. Because all of the fundamental fluid mechanics are computed, 

the code is capable of capturing the nonlinear characteristics of the flow fields of interest. With the actual 

modeling of the grid movement of the blade rows in relative motion, this code is capable of computing 

the unsteady interactions between blade rows. Details of the flow solver are given by Chen and Whitfield

(Ref. 14). The approach to parallelization for large-scale, complex problems is discussed by Chen and 

Briley (Ref. 15). TURBO has been previously modified for application to study flutter and forced 

vibration (Refs. 16 to 18). TURBO has been further validated in several previous research efforts

regarding aeromechanics (Refs. 19 to 24).

Computational Research Test Article

A high-speed, forward swept fan has been developed for research purposes. Its raison d’être was the 

study of new design concepts with respect to noise and acoustics. In the course of development and 

testing, aeromechanical concerns arose. The fan was aeroelastically stable in most operating conditions, 

but flutter was observed during physical testing at low-flow conditions above the operating line.

Concomitant and complementary CFD analysis using TURBO proved to accurately simulate the 

aeroelastic stability (and instability) of the fan for the mode of concern along the speedline of concern.

Given the corroboration and mutual validation of physical experiment with this fan and TURBO 

simulations of this fan, this fan is ripe for studying aeroelastic response to inlet distortion. Will previously 

stable points become unstable? Will already-unstable points become more unstable? Will previouslyunstable points become stable? Also of interest, how will fan aerothermodynamic performance be 

impacted?

Physical and Computational Definition of Fan

The fan of interest is comprised of twenty-two blades. The fan was designed for purposes of noise 

reduction in a typical podded-engine mounting. The fan was never intended to be part of an embedded 

propulsion system with distorted inlet flow, but it is studied in this vein because of corroboration between 

the computational and experimental results with regard to flutter assessment (with clean inlet flow).

Computationally, the fan is preceded by inlet duct of length-ratio 1.373 with respect to blade axial chord 

and followed by exit duct of length-ratio 2.364 with respect to blade axial chord. Physically, the average 

radial tip gap at the rotational speed of interest is about 98.4 percent of blade radial span; in the initial 

study, four computational volumes filled this tip gap. A CFD H-mesh representation of the fan is pictured 

in Figure 1; this refined mesh was constructed in the course of this study.

Initial Simulation and Validation

In the initial application of TURBO for the aeroelastic study of this fan, single-passage simulations 

with appropriate periodic or phase-lag boundary conditions were employed. With the desire to study inlet 

distortion typical to serpentine ducts and embedded inlets, single-passage and even multi-passage 

periodic-sector models are not adequate. Full-annulus simulations must be performed, at the (large) added 

expense of computational resource requirements. In-house codes used to pre-process TURBO inputs and 

post-process TURBO outputs have been developed and modified to facilitate full-annulus TURBO 

simulations for both aerodynamic and aeroelastic studies. Figures follow which demonstrate the 

consistency of solution from serial TURBO single-passage solution to parallel TURBO full-annulus 

solution of the original grid. Each of these charts arises from clean inlet (axisymmetric radial distortion) 
NASA/TM—2011-216959 4

flows. The general aeroelastic response of the fan has been retained as shown in Figure 2, while slight 

circumferential blade-to-blade variations become more pronounced in higher nodal-diameter modes of the 

full annulus configuration in Figure 3. Such circumferential variation in response to clean inlet flow 

further incites the study which follows.

 Figure 1.—(Refined) H-Mesh of research fan.

Figure 2.—Consistent results through code evolution and solution 

refinement. (Peak efficiency, original grid, static pressure profile 

exit)

-12 -9 -6 -3 0 3 6 9 12 Aerodynamic Damping

Nodal Diameter

Single Passage Time-Shift Full Annulus
NASA/TM—2011-216959 5

Figure 3.—Slight blade-to-blade variation observed in full-annulus 

simulations. (Peak efficiency, original grid, static pressure profile exit.) 

Computational Considerations

Refining the Grid

The original grid proved exceptional in computational efficiency for solution of single-passage 

configurations (periodic and time-shift alike) with constant static pressure profile (“SPP”) exit boundary 

conditions. However, the nature of inlet distortion suggests that a throttle exit boundary condition, which 

would not prescribe nor enforce two-dimensional spatial distribution of any aerothermodynamic 

quantities within the exit plane, facilitates better solution to such flow conditions than a constant static 

pressure profile exit boundary condition. A throttle exit boundary condition is thus applied to this original 

grid. Figure 4 shows the excellent overlap between simulations with constant static pressure profile exits 

and simulations with throttle exits (“ECMF”: Exit Corrected Mass Flow) from choke through peakefficiency and slightly beyond. Approaching stall, however, the throttle does not adequately converge to a 

single, quotable result for this original grid. Previous computational research using TURBO on nonaxisymmetric flows documents TURBO’s success in simulating nonaxisymmetric and/or unstable flows 

with the throttle boundary condition (Ref. 25). Attributing the failure of the throttle boundary condition 

with this original grid to mesh density, the author refines the grid with 25 percent more volumes spanwise

in the tip gap, 20 percent more volumes spanwise on the blade, 41 percent more volumes streamwise on 

the blade, 180 percent additional streamwise volumes in the inlet, and 156 percent additional streamwise 

volumes in the exit grid. The significant refinement in the streamwise direction was deemed critical in 

preserving the non-axisymmetric character of the flow through its streamwise progress. The refined grid 

also contains 16 percent additional volumes circumferentially. The demand for computational resources – 

directly proportional to total volume count – was a practical consideration throughout the refinement 

process.

1 4 7 10 13 16 19 22

Aerodynamic Damping

Blade Index

0ND 2NDF 4NDF 6NDF

11ND 6NDB 4NDB 2NDB
NASA/TM—2011-216959 6

 Figure 4.—Throttle exit boundary condition does not converge with full flow 

range of constant static pressure profile exit boundary condition on 

original grid.

Inlet Distortion Boundary Condition

The crux of this study lies in accurately capturing and maintaining a prescribed inlet distortion 

boundary condition. TURBO has long supported an isentropic inlet boundary condition (Ref. 14) wherein 

total conditions (with variation allowed only in the radial direction) are held as dictated by the 

characteristics. This boundary condition is very well suited for stable flows which exhibit variation 

between hub and casing (axisymmetrically) only, but this boundary condition must be modified to allow 

for circumferential variation when studying inlet distortion flows. For this study, the inlet boundary 

condition is modified to allow for user-specified prescription of total temperature, total pressure, radial 

flow angle, and tangential (“swirl”) flow angle about a user-specified orthogonal polar grid. To enforce 

the prescribed circumferentially-varying flow conditions before a rotor, the distortion pattern must be 

interpolated, at each time step, upon backward-rotating rotor grids. (N.B.: TURBO computes the passage 

flows using stationary grids and applies appropriate body forces.)

Of particular interest to NASA in the development of propulsion systems for the hybrid wing/body 

aircraft concept is the BLI inlet concept, studied by Berrier et al. (Ref. 26) at Langley Research Center, 

known as “Inlet A”. This inlet’s total pressure (PT) distortion pattern, approximated as planar-symmetric, 

is shown in Figure 5. For comparison purposes, the clean inlet distribution of inlet total pressure for 

which this fan was designed is shown in Figure 6. The Inlet A pattern features a 10.5 percent spread in PT, 

while the lowest PT in the clean inlet is 12 percent less than the clean inlet maximum PT. The areaaveraged total pressure of the distorted inlet is about 4 percent less than the area-averaged total pressure 

of the clean inlet. No circumferential distortion in total temperature, radial flow angle, or swirl is 

incorporated in the present study due to lack of data. Pressure Ratio, Total to Static

Physical Mass Flow

ECMF SPP
NASA/TM—2011-216959 7

Figure 5.—Approximation of Inlet A distortion pattern 

(total pressure; red is high). Figure 6.—Design clean inlet total pressure distribution 

(red is high).

Validating the Throttle Exit Boundary Condition for This Study

Having refined the grid to enhance the accuracy and applicability of the throttle exit boundary 

condition for nonaxisymmetric flow simulations such as these, the behavior of this exit boundary 

condition on the refined grid must be verified. In Figure 7, the behaviors of exit constant static pressure 

profile, exit constant static pressure with radial equilibrium (“PRE”) held at hub, and exit corrected mass 

flow boundary conditions are plotted. As shown in Figure 7, with appropriate refinement, the exit 

corrected mass flow boundary condition emulates the aerothermodynamic performance of the exit 

constant static pressure profile and exit constant static pressure with radial equilibrium held at the hub, 

while this throttle boundary condition also demonstrates (more accurate) increased range from choke to 

stall.

Given the impetus to employ this boundary condition, it is further instructive to inspect the resulting 

exit pressure field when unconstrained with regard to spatial distribution. Figure 8 depicts the

(expectedly) very axisymmetric distribution of exit static pressure from a full annulus simulation of the 

refined grid with the throttle exit boundary condition and a clean inlet (cf Figure 6). Next in Figure 9, the 

exit static pressure field from a full annulus simulation of the refined grid with the throttle exit boundary 

condition and the distorted inlet (cf Fig. 5) is shown. Clearly, non-axisymmetric and planar asymmetric 

flow conditions remain at the exit plane, and thus an exit boundary condition specifying uniformity, 

axisymmetry, circumferential periodicity, or planar symmetry would be (inaccurately) over-constraining 

the exit flow field. Thus the corrected mass flow “throttle” exit boundary condition is deemed most 

appropriate for further study of “Inlet A” inlet distortion flows.
NASA/TM—2011-216959 8

Figure 7.—Throttle exit boundary condition exhibits greater flow range 

than static pressure profile and static pressure/radial equilibrium

boundary conditions on refined grid.

Figure 8.—Exit static pressure distribution from clean 

inlet simulation (Red is high).

Figure 9.—Exit static pressure distribution from distorted 

inlet simulation (Red is high).

Clean Inlet v Distorted Inlet: Aerothermodynamic Performance

As is readily understood and previously demonstrated here, the “local” aerodynamic behavior of the 

fan is greatly impacted by the condition of the inlet flow. The global aerothermodynamic performance is 

also impacted by inlet distortion: Figure 10 demonstrates the decreased physical flow at choke for the 

distorted inlet scenario: the 4 percent decrement in physical mass flow choke is directly commensurate 

with the 4 percent decrement in area-averaged inlet total pressure. Of greater interest, however, is the 

overall performance as gauged by pressure ratio versus corrected flow; see Figure 11. Observe that when 

viewed from the corrected basis, the pressure recovery is nearly unchanged between the clean inlet and 

the distorted (weakened) inlet. The pumping behavior of the compression system is almost independent of 

inlet flow condition, whether it be a “global”, “absolute” change in inlet pressure (e.g., flight altitude) or 

the “local”, “relative” change in inlet pressure inherent to inlet distortion. Presure Ratio, Total-to-Static

Physical Mass Flow

SPP PRE ECMF
NASA/TM—2011-216959 9

Figure 10.—Pressure ratio v physical mass flow for clean and distorted inlets.

Figure 11.—Pressure ratio v corrected mass flow for clean and distorted inlets.

Clean Inlet v Distorted Inlet: Aeroelastic Response

The primary research interest here is the impact of a BLI inlet’s distorted inflow on aeroelastic 

response of a fan. The subject test article has been shown (computationally and experimentally) to be 

aeromechanically stable near peak-efficiency for all blade-disk modes of interest with a clean inlet. 

However, experiment reveals aeromechanical instability as the machine approaches stall with a clean 

inlet. Previous TURBO analysis (single passage, relatively coarse grid, constant exit static pressure 

profile; not shown here) reveals a minimum of aeromechanical stability for the same mode and flow 

condition which had been observed as unstable in the physical experiment. Pressure Ratio, Total-to-Static

Physical Mass Flow 

Inlet A

Clean

Pressure Ratio, Total-to-Static

Corrected Mass Flow 

Inlet A

Clean
NASA/TM—2011-216959 10

Figure 12.—Aeroelastic sweeps for clean inlet and Inlet A, peak-efficiency 

and near-stall. Refined grid, full-annulus, throttle exit boundary condition.

Figure 12 plots the aeroelastic sweeps (full annulus, refined grid, throttle exit) for a first flexural 

mode about the nodal diameters of interest for both clean and distorted inlets at both peak-efficiency and 

near-stall flow conditions. The expanded (full annulus), refined grid with the throttle exit indeed shows 

the two nodal diameter forward traveling wave mode to be unstable, rather than minimally stable, with 

clean inlet flow at the near-stall condition. For both peak-efficiency and near-stall flow conditions, the 

aeroelastic stability for all nodal diameter (ND) wave modes is minimally changed when viewing the 

distorted inlet flows from an annular-averaged perspective. The Inlet A near-stall simulation of the 2NDB 

wave and the 4NDF wave modes did not satisfactorily converge in the time frame of this paper. 

All aerodynamic damping quantities in Figure 12 represent annular averages of the twenty-two blades 

about the annulus. Observing that the magnitudes and general shapes of the aeroelastic sweeps are similar 

among all conditions (peak-efficiency and near-stall, clean inlet and Inlet A), the premise of averaging 

about the annulus seems reasonable, but closer inspection of the data, temporally and spatially, is quite 

revealing. In Figure 13, the fan has a consistent aeroelastic response to clean inlet flow near stall. The 

aerodynamic work of both the 2NDF and 11ND modes oscillates with small amplitude about the annulus, 

but it is clearly evident that the 11ND mode is stable while the 2NDF mode is unstable as is shown in 

Figure 12. Figure 14 plots the convergence of the annular-averaged aerodynamic work per cycle for the 

11ND and 2NDF modes behind clean and distorted inlets; all converge nicely in their aeroelastic 

responses, irrespective of clean or distorted inlet condition, though the 2NDF modes are slightly unstable 

while the 11ND modes are solidly stable.

Figures 15 and 16 focus exclusively on the distorted inlet condition. Figure 15 shows the 180° phaseshift in vibratory motion between neighboring blades for the 11ND mode, while Figure 16 clearly shows 

the in-phase vibratory motion of blades spatially separated by 180° about the annulus for the 2NDF mode. 

In both figures, the aerodynamic work is plotted at the angular position of the blade at the completion of 

the vibratory cycle. The angular positions of the annulus’ highest and lowest total pressures are annotated 

along with the direction of blade rotation to clarify the range of pressures through which the blade swept 

in the course of the single respective vibratory cycle. In Figure 15, the minimally-oscillating clean-inlet 

aerodynamic work is plotted; it is readily apparent that the annular-averaged aerodynamic work differs 

little between distorted inlet and clean inlet; the clean inlet is solidly stable. The borderline-instability of 

the 2NDF mode is demonstrated in Figure 16 by the slightly-positive clean-inlet aerodynamic work which 

-12 -9 -6 -3 0 3 6 9 12 Aerodynamic Damping 

Nodal Diameter

Clean PE

Inlet A PE

Clean NS

Inlet A NS

Stable

Unstable
NASA/TM—2011-216959 11

again coincides with the mean of the distorted inlet’s aerodynamic work variation about the annulus. In 

addition to its loss of global stability, the 2NDF mode also features a much greater band of oscillation 

(variation by angular position)—about 60 times the magnitude of its mean—in its aeroelastic response 

than the 11ND mode, which oscillates with a bandwidth of 10 times its mean.

Figure 13.—Aerodynamic work v angular position, near-stall, 11ND and 

2NDF, clean inlet, all blades.

Figure 14.—Annular-averaged aerodynamic work v vibration cycle, nearstall.

-2.0E-07

-1.5E-07

-1.0E-07

-5.0E-08

0.0E+00

5.0E-08

0 30 60 90 120 150 180 210 240 270 300 330 360

Aerodynamic Work

Angular Position

at end of Vibration Cycle, deg

Clean 2NDF Clean 11ND

Unstable

Stable

-1.0E-07

-8.0E-08

-6.0E-08

-4.0E-08

-2.0E-08

0.0E+00

2.0E-08

0 1 2 3 4 5 6 7 8 9

Aerodynamic Work

Vibration Cycle

Clean, 11ND Clean, 2NDF Inlet A, 11ND Inlet A, 2NDF

Unstable

Stable
NASA/TM—2011-216959 12

Figure 15.—Aerodynamic work v angular position, near-stall, 11ND, Inlet 

A, all blades.

Figure 16.—Aerodynamic work v angular position, near-stall, 2NDF, Inlet 

A, blades 1 and 12.

Conclusions and Recommendations

An established CFD code, validated in the analysis of aerothermodynamic performance, fluid 

dynamics, and aeroelastics, has been modified and applied in a preliminary study of the effects of inlet 

distortion on aeroelastic stability of a high-speed forward-swept fan. Care was taken throughout the 

modifications of the code to ensure that numerical accuracy of the code was maintained. In developing 

the CFD representation of the fan, care was taken to ensure adequate gridding of the domain was 

prescribed for reliable solutions and that the simulations were manageable within available computational 

resources. The existing axisymmetric “total condition preserved” inlet boundary condition has been 

modified to handle circumferential nonuniformity.

-4.50E-07

-3.00E-07

-1.50E-07

0.00E+00

1.50E-07

3.00E-07

0 30 60 90 120 150 180 210 240 270 300 330 360

Aerodynamic Work

Angular Position

at end of Vibration Cycle, deg

Clean Inlet

Rotation

High PT

Low PT

Negative: Stable

Positive: 

Unstable

-6.00E-07

-4.50E-07

-3.00E-07

-1.50E-07

0.00E+00

1.50E-07

3.00E-07

4.50E-07

6.00E-07

7.50E-07

0 30 60 90 120 150 180 210 240 270 300 330 360

Aerodynamic Work

Angular Position

at end of Vibration Cycle, deg

Blade 1 Blade 12 Clean Inlet

Rotation

High PT

Positive: Unstable Low PT

Negative: 

Stable
NASA/TM—2011-216959 13

Application of the throttle exit boundary condition, allowing the exit flow to attain equilibrium

without artificial constraint, has shown that inlet distortions do not fully attenuate upon passage through 

single rotor. The throttle boundary condition effectively emulated other static pressure exit boundary 

conditions in terms of global aerothermodynamic performance data for the given grid. It was observed 

when applying an inlet distortion of the nature of Inlet A to this research fan that distorted inlet 

aerothermodynamic performance was very similar to clean inlet performance when plotted on a corrected 

flow basis.

When examining aeroelastic behavior, the clean inlet solutions yield much cleaner, consistent 

solutions. The distorted inlet aeroelastic simulations introduce great oscillation to solidly stable flow 

conditions like the 11ND near-stall, Inlet A simulation; the aeroelastic response quantities oscillate much 

more greatly for the 2NDF near-stall Inlet A simulation. While the initial investigation into the effects of 

inlet distortion on aerodynamic damping suggest that the aeroelastic stability of the fan is minimally

impacted by the inlet distortion, the magnitudes of the oscillations within the distorted flow simulations 

encourage further investigation into the details of the flow field. Much detailed numerical data has been 

produced and archived through the course of this study, and there is much left to investigate. It will be of 

great interest to interrogate the data in an unsteady manner. Each blade’s aeroelastic response may be 

traced through the blade’s temporal and spatial progression; some preliminary data has been presented 

here. In-depth study of this data may provide insight as to the dominant/critical issues of inlet distortion 

which have great adverse impact on aeromechanical stability and which factors/conditions present 

minimal adverse impact on aeromechanical stability.

References

1. Liebeck, R., “Design of the Blended-Wing-Body Subsonic Transport,” AIAA Journal of Aircraft,

Vol. 41, No. 1, pp. 10–25, 2004.

2. Smith, Leroy H., Jr., “Wake Ingestion Propulsion Benefit,” AIAA Journal of Propulsion and Power, 

Vol. 9, No. 1, pp. 74–82, 1993.

3. Plas, A., Crichton, D., Sargeant, M., Hynes, T., Greitzer, E., Hall, C., and Madani, V., “Performance 

of a Boundary Layer Ingesting (BLI) Propulsion System,” AIAA–2007–0450, Jan. 2007.

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4. TITLE AND SUBTITLE

Effects of Inlet Distortion on Aeromechanical Stability of a Forward-Swept High-Speed Fan 

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Herrick, Gregory, P. 

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14. ABSTRACT

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are 

partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic 

performance of the compression system in response to inlet distortion. Separately, a forward-swept high-speed fan was developed to address 

noise concerns of modern podded turbofans; however this fan encounters aeroelastic instability (flutter) as it approaches stall. A threedimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean 

inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions, is 

modified and then applied in a computational study to preliminarily assess the effects of inlet distortion on aeroelastic stability of the fan. 

Computational engineering application and implementation issues are discussed, followed by an investigation into the aeroelastic behavior 

of the fan with clean and distorted inlets. 

15. SUBJECT TERMS

Aeroelasticity; Turbomachinery; Turbofan; Flutter; Computational fluid dynamics; Inlet distortion; Vibration 

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