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

2010 ◽  
Author(s):  
Gregory Herrick
Keyword(s):  
High Speed ◽  
Inlet Distortion ◽  
Aeromechanical Stability

Blade Forcing Function and Aerodynamic Work Measurements in a High Speed Centrifugal Compressor With Inlet Distortion

2010 ◽  
Vol 132 (9) ◽  
Author(s):  
Albert Kammerer ◽  
Reza S. Abhari
Keyword(s):  
High Speed ◽  
Pressure Sensors ◽  
Forced Response ◽  
Blade Surface ◽  
Centrifugal Compressors ◽  
Flow Distortion ◽  
Unsteady Pressure ◽  
Inlet Distortion ◽  
Forcing Function ◽  
Work Distribution

Centrifugal compressors operating at varying rotational speeds, such as in helicopters or turbochargers, can experience forced response failure modes. The response of the compressors can be triggered by aerodynamic flow nonuniformities such as with diffuser-impeller interaction or with inlet distortions. The work presented here addresses experimental investigations of forced response in centrifugal compressors with inlet distortions. This research is part of an ongoing effort to develop related experimental techniques and to provide data for validation of computational tools. In this work, measurements of blade surface pressure and aerodynamic work distribution were addressed. A series of pressure sensors were designed and installed on rotating impeller blades and simultaneous measurements with blade-mounted strain gauges were performed under engine representative conditions. To the best knowledge of the authors, this is the first publication, which presents comprehensive experimental unsteady pressure measurements during forced response, for high-speed radial compressors. The experimental data were obtained for both resonance and off-resonance conditions with uniquely tailored inlet distortion. This paper covers aspects relating to the design of fast response pressure sensors and their installation on thin impeller blades. Additionally, sensor properties are outlined with a focus on calibration and measurement uncertainty estimations. The second part of this paper presents unsteady pressure results taken for a number of inlet distortion cases. It will be shown that the intended excitation order due to inlet flow distortion is of comparable magnitude to the second and third harmonics, which are consistently observed in all measurements. Finally, an experimental method will be outlined that enables the measurement of aerodynamic work on the blade surface during resonant crossing. This approach quantifies the energy exchange between the blade and the flow in terms of cyclic work along the blade surface. The phase angle between the unsteady pressure and the blade movement will be shown to determine the direction of energy transfer.

Shock Propagation and MPT Noise From a Transonic Rotor in Nonuniform Flow

2012 ◽  
Vol 135 (1) ◽  
Author(s):  
Jeffrey J. Defoe ◽  
Zoltán S. Spakovszky
Keyword(s):  
High Speed ◽  
Acoustic Energy ◽  
Shock Propagation ◽  
Far Field ◽  
Nonuniform Flow ◽  
Mean Flow ◽  
Flow Distortion ◽  
Inlet Distortion ◽  
Circumferential Extent ◽  
Transonic Rotor

One of the major challenges in high-speed fan stages used in compact, embedded propulsion systems is inlet distortion noise. A body-force-based approach for the prediction of multiple-pure-tone (MPT) noise was previously introduced and validated. In this paper, it is employed with the objective of quantifying the effects of nonuniform flow on the generation and propagation of MPT noise. First-of-their-kind back-to-back coupled aero-acoustic computations were carried out using the new approach for conventional and serpentine inlets. Both inlets delivered flow to the same NASA/GE R4 fan rotor at equal corrected mass flow rates. Although the source strength at the fan is increased by 38 dB in sound power level due to the nonuniform inflow, far-field noise for the serpentine inlet duct is increased on average by only 3.1 dBA overall sound pressure level in the forward arc. This is due to the redistribution of acoustic energy to frequencies below 11 times the shaft frequency and the apparent cut-off of tones at higher frequencies including blade-passing tones. The circumferential extent of the inlet swirl distortion at the fan was found to be two blade pitches, or 1/11th of the circumference, suggesting a relationship between the circumferential extent of the inlet distortion and the apparent cut-off frequency perceived in the far field. A first-principles-based model of the generation of shock waves from a transonic rotor in nonuniform flow showed that the effects of nonuniform flow on acoustic wave propagation, which cannot be captured by the simplified model, are more dominant than those of inlet flow distortion on source noise. It demonstrated that nonlinear, coupled aerodynamic and aero-acoustic computations, such as those presented in this paper, are necessary to assess the propagation through nonuniform mean flow. A parametric study of serpentine inlet designs is underway to quantify these propagation effects.

scholarly journals Stall Inception in a 5-Stage HP-Compressor With Increased Load due to Inlet Distortion

1999 ◽  
Author(s):  
Werner Jahnen ◽  
Thomas Peters ◽  
Leonhard Fottner
Keyword(s):  
Linear Stability ◽  
Test Data ◽  
Total Pressure ◽  
High Speed ◽  
Hot Wire ◽  
Stall Inception ◽  
Diffusion Factor ◽  
Inlet Distortion ◽  
And Performance ◽  
And Diffusion

Unsteady measurements of the flow and performance of a high speed 5-stage HP compressor have been carried out at different speeds under undistorted conditions and with swirl and total pressure inlet distortions. Distributions of incidence and diffusion factor have been derived from the test data which, together with hot-wire measurements of stall inception, provide new insights into the onset of stall with inlet distortion. A stall cell initiates as a disturbance in the distorted flow sector, which may decay as it passes through the undistorted sector. Stall inception occurs only when the damping of the disturbance in the undistorted sector is insufficient to prevail its growth. As this damping depends on the size of the disturbance, the Parallel Compressor model, based on the linear stability properties of the undistorted compressor alone, is unable to predict the stall inception with inlet distortion.

scholarly journals Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part I — Radial Distortion

1998 ◽  
Author(s):  
Z. S. Spakovszky ◽  
H. J. Weigl ◽  
J. D. Paduano ◽  
C. M. van Schalkwyk ◽  
K. L. Suder ◽  
...  
Keyword(s):  
Mass Flow ◽  
High Speed ◽  
Transfer Functions ◽  
Axial Flow ◽  
Rotating Stall ◽  
Forced Response ◽  
Single Stage ◽  
Nasa Lewis Research ◽  
Radial Distortion ◽  
Inlet Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An annular array of 12 jet-injectors located upstream of the rotor tip was used for forced response testing and to extend the compressor stable operating range. Results for radial distortion are reported in this paper. First, the effects of radial distortion on the compressor performance and the dynamic behavior were investigated. Control laws were designed using empirical transfer function estimates determined from forced response results. The transfer functions indicated that the compressor dynamics are decoupled with radial inlet distortion, as they are for the case of undistorted inlet flow. Single-input-single-output (SISO) control strategies were therefore used for the radial distortion controller designs. Steady axisymmetric injection of 4% of the compressor mass flow resulted in a reduction in stalling mass flow of 9.7% relative to the case with inlet distortion and no injection. Use of a robust H∞ controller with unsteady non-axisymmetric injection achieved a further reduction in stalling mass flow of 7.5%, resulting in a total reduction of 17.2%.

Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part I—Radial Distortion

1999 ◽  
Vol 121 (3) ◽  
pp. 510-516 ◽  
Author(s):  
Z. S. Spakovszky ◽  
H. J. Weigl ◽  
J. D. Paduano ◽  
C. M. van Schalkwyk ◽  
K. L. Suder ◽  
...  
Keyword(s):  
Mass Flow ◽  
High Speed ◽  
Transfer Functions ◽  
Axial Flow ◽  
Rotating Stall ◽  
Forced Response ◽  
Single Stage ◽  
Nasa Lewis Research ◽  
Radial Distortion ◽  
Inlet Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An annular array of 12 jet-injectors located upstream of the rotor tip was used for forced response testing and to extend the compressor stable operating range. Results for radial distortion are reported in this paper. First, the effects of radial distortion on the compressor performance and the dynamic behavior were investigated. Control laws were designed using empirical transfer function estimates determined from forced response results. The transfer functions indicated that the compressor dynamics are decoupled with radial inlet distortion, as they are for the case of undistorted inlet flow. Single-input-single-output (SISO) control strategies were therefore used for the radial distortion controller designs. Steady axisymmetric injection of 4 percent of the compressor mass flow resulted in a reduction in stalling mass flow of 9.7 percent relative to the case with inlet distortion and no injection. Use of a robust H∞ controller with unsteady nonaxisymmetric injection achieved a further reduction in stalling mass flow of 7.5 percent, resulting in a total reduction of 17.2 percent.

Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part II—Circumferential Distortion

1999 ◽  
Vol 121 (3) ◽  
pp. 517-524 ◽  
Author(s):  
Z. S. Spakovszky ◽  
C. M. van Schalkwyk ◽  
H. J. Weigl ◽  
J. D. Paduano ◽  
K. L. Suder ◽  
...  
Keyword(s):  
Mass Flow ◽  
High Speed ◽  
Axial Flow ◽  
Rotating Stall ◽  
Forced Response ◽  
Feedback Stabilization ◽  
Single Stage ◽  
Nasa Lewis Research ◽  
Control Laws ◽  
Inlet Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An array of 12 jet injectors located upstream of the compressor was used for forced response testing and feedback stabilization. Results for a circumferential total pressure distortion of about one dynamic head and a 120 deg extent (DC(60) = 0.61) are reported in this paper. Part I (Spakovszky et al., 1999) reports results for radial distortion. Control laws were designed using empirical transfer function estimates determined from forced response results. Distortion introduces coupling between the harmonics of circumferential pressure perturbations, requiring multivariable identification and control design techniques. The compressor response displayed a strong first spatial harmonic, dominated by the well-known incompressible Moore–Greitzer mode. Steady axisymmetric injection of 4 percent of the compressor mass flow resulted in a 6.2 percent reduction of stalling mass flow. Constant gain feedback, using unsteady asymmetric injection, yielded a further range extension of 9 percent. A more sophisticated robust H∞ controller allowed a reduction in stalling mass flow of 10.2 percent relative to steady injection, yielding a total reduction in stalling mass flow of 16.4 percent.

An Argument for Enhancement of the Current Inlet Distortion Ground Test Practice for Aircraft Gas Turbine Engines

2001 ◽  
Author(s):  
Milt Davis ◽  
Alan Hale ◽  
Dave Beale
Keyword(s):  
Total Pressure ◽  
High Speed ◽  
High Performance ◽  
Numerical Study ◽  
Stability Limit ◽  
Interface Plane ◽  
Turbine Engine ◽  
Inlet Distortion ◽  
Current Test ◽  
Pressure Profiles

The current high-performance aircraft development programs, and the trends in research and development activities suggest a rapidly increasing level of aircraft subsystem integration, particularly between the airframe/inlet and the propulsion system. Traditionally these subsystems have been designed, analyzed, and tested as isolated systems. The interaction between the subsystems is modeled primarily through evaluating inlet distortion in an inlet test and simulating this distortion in engine tests via screens or similar devices. For the current test methodology, the environment that is supplied by the inlet is simulated by the imposition of total pressure profiles at the aerodynamic interface plane (AIP). Unsteady or transient variation in total pressure is generally not considered to be important. In addition, angular flow, commonly called swirl, is also not considered important enough to be simulated. In the current paper, an overview of current techniques for inlet performance, distortion characterization, and engine distortion testing is presented. A numerical study was conducted on a single high-speed rotor to qualify potential effects on stability and performance and to support the concept that dynamic distortion and swirl may have large enough effects to affect the experimentally determined stability limit. This paper reports a numerical investigation using a 3D compression system simulation that supports the enhancement of the existing methodology to include the effects of time-dependent distortion and swirl effects. Based upon both experimental and numerical evidence, AEDC has embarked on efforts to develop inlet simulator technologies directed toward future airframe-propulsion integration requirements. This paper presents issues that require advancements in the simulation of inlet distortion techniques for direct-connect turbine engine tests.

scholarly journals Rotating Stall Control in a High-Speed Stage With Inlet Distortion: Part II — Circumferential Distortion

1998 ◽  
Author(s):  
Z. S. Spakovszky ◽  
C. M. van Schalkwyk ◽  
H. J. Weigl ◽  
J. D. Paduano ◽  
K. L. Suder ◽  
...  
Keyword(s):  
Mass Flow ◽  
High Speed ◽  
Axial Flow ◽  
Rotating Stall ◽  
Forced Response ◽  
Feedback Stabilization ◽  
Single Stage ◽  
Nasa Lewis Research ◽  
Control Laws ◽  
Inlet Distortion

This paper presents the first attempt to stabilize rotating stall in a single-stage transonic axial flow compressor with inlet distortion using active feedback control. The experiments were conducted at the NASA Lewis Research Center on a single-stage transonic core compressor inlet stage. An array of 12 jet injectors located upstream of the compressor was used for forced response testing and feedback stabilization. Results for a circumferential total pressure distortion of about one dynamic head and a 120° extent (DC(60) = 0.61) are reported in this paper. Part I (Spakovszky et al. (1998)) reports results for radial distortion. Control laws were designed using empirical transfer function estimates determined from forced response results. Distortion introduces coupling between the harmonics of circumferential pressure perturbations, requiring multi-variable identification and control design techniques. The compressor response displayed a strong first spatial harmonic, dominated by the well known incompressible Moore-Greitzer mode. Steady axisymmetric injection of 4% of the compressor mass flow resulted in a 6.2% reduction of stalling mass flow. Constant gain feedback, using unsteady asymmetric injection, yielded a further range extension of 9%. A more sophisticated robust controller allowed a reduction in stalling mass flow of 10.2% relative to steady injection, yielding a total reduction in stalling mass flow of 16.4%.

Inlet Distortion Simulations of a Transonic Fan With CFD and Low-Order Methods

2021 ◽  
Author(s):  
Carlo Favaron ◽  
Andrea Magrini ◽  
Alessandro Visentin ◽  
Luca Menegozzo ◽  
Ernesto Benini
Keyword(s):  
High Speed ◽  
Similar Response ◽  
Inlet Distortion ◽  
Speed Flow ◽  
The Mean ◽  
High Level ◽  
Transonic Fan ◽  
Operating Points ◽  
Inflow Conditions ◽  
Low Order

Abstract Future aircraft are expected to have a high level of fuselage-engine integration, exposing the propulsor to non-uniform inflow conditions, in which the performance and stability of the engine fan can be severly affected. This paper proposes a study of a transonic fan subject to inlet distortion by employing steady RANS simulations and mean line calculations. The steady CFD results, although insufficient to correctly measure the interaction of the distortion across the cascade passages, are used as a benchmark to compare the prediction of a leaner low-order tool, adopting a superposition of several 1D calculations in the discretised compressor approach. A continuous total pressure deficit is imposed at the inlet and the distorted operating points are compared to the clean inflow case. The performance drop at the peak efficiency point is similar for the two models, altough the mean-line 1D method fails in closely matching the distribution in the upper span of the blade. The qualitatively similar response of the low-order approach to the distorted inflow should allow for its use in preliminary blade design exploration once improved for high-speed flow prediction.

A Method for Evaluating the Effect of Circumferential Inlet Distortion on the Aerodynamic Stability of Multi-Stage Axial-Flow Compressors

2010 ◽  
Author(s):  
Tsuguji Nakano ◽  
Andy Breeze-Stringfellow
Keyword(s):  
Steady State ◽  
High Speed ◽  
Axial Flow ◽  
Pressure Rise ◽  
Stability Limit ◽  
Rotating Stall ◽  
Classical Dynamic ◽  
Inlet Flow ◽  
Inlet Distortion ◽  
Multi Stage

A simple engineering parameter to evaluate the stability of high-speed multi-stage compressors with distorted inlet flow has been derived based on a simplified semi-compressible linear stability model. The parameter consists of steady-state flow quantities and geometric parameters of the compressor and it indicates that the circumferential integral of the slope of the steady-state individual blade row static pressure rise characteristics is important in the determination of the compressor stability limit in the presence of distortion. The parameter reduces to the author’s rotating stall inception parameter in the limit of non-distorted inlet flow. Since the model includes a downstream plenum and throttle, a condition for pure surge inception with undistorted inlet flow has been deduced. The pure surge conditions can be reduced to the classical dynamic and static instability conditions in the limit of a constant annulus area incompressible compressor. The results indicate that rotating stall always precedes surge instability, as many engineers and researchers would expect from experience. The parameter for instability with inlet distortion was calculated using test data measured in a high-speed 5-stage compressor with two different types of circumferential inlet distortion, and the results show that the parameter has a strong correlation with the data and is an improvement over the classical incompressible stability parameter. The results demonstrate that the parameter captures much of the physics important during the instability inception in a high-speed multi-stage compressor subjected to circumferential inlet distortion. The parameter clearly shows how each compressor component’s characteristics contribute to the overall stability in a high speed axial multi-stage compressor, therefore, it will aid engineers and designers in their understanding and prediction of the aerodynamic instability inception phenomena.

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