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NASA NTRS1999

Inlet Flow Distortion and Unsteady Blade Response in a Transonic Axial-Compressor Rotor

This paper describes the unsteady blade surface pressures on the first-stage rotor blades of a two-stage transonic axial flow compressor experiencing inlet flow distortion. This study was conducted to demonstrate the ability of a full annulus unsteady Reynolds-averaged Navier-Stokes numerical technique to predict unsteady pressures on the rotor blades operating in a distorted inflow. A total pressure distortion produced by a variable mesh screen mounted near the inlet was used to excite the unsteady blade loading on the rotor. On-blade pressure transducers were used to measure the unsteady blade surface pressure. These pressures and the resulting transient load on the rotor blades were compared to the numerical prediction. It is important to develop numerical techniques to predict these transient loads to better understand the response of compressor blades to forcing functions. With this enhanced understanding and ability to predict these transient forces, more robust compressors can be developed. In the study, a high definition of the inlet flow distortion was achieved by rotating the distortion screens. In this manner the inlet flow distortion and the distortion at the first stage stator leading edge were measured at approximately every 0.7 degrees. This full annulus high definition of the inlet flow distortion was used as the inlet boundary condition for the numerical technique. The experimental measurements and numerical analyses are highly complementary in this study. Detailed comparisons between the measurements and the numerical analyses indicate that the current numerical procedure calculates the unsteady aerodynamic pressure on the blade surfaces reasonably well. Further, the agreement of the measured and predicted rotor exit flow distortion at the first stage stator leading edge provides verification of the numerical technique.

Rabe, D. C.↗
NASA NTRS2001

A Robust Design Methodology for Optimal Microscale Secondary Flow Control in Compact Inlet Diffusers

It is the purpose of this study to develop an economical Robust design methodology for microscale secondary flow control in compact inlet diffusers. To illustrate the potential of economical Robust Design methodology, two different mission strategies were considered for the subject inlet, namely Maximum Performance and Maximum HCF Life Expectancy. The Maximum Performance mission maximized total pressure recovery while the Maximum HCF Life Expectancy mission minimized the mean of the first five Fourier harmonic amplitudes, i.e., 'collectively' reduced all the harmonic 1/2 amplitudes of engine face distortion. Each of the mission strategies was subject to a low engine face distortion constraint, i.e., DC60<0.10, which is a level acceptable for commercial engines. For each of these missions strategies, an 'Optimal Robust' (open loop control) and an 'Optimal Adaptive' (closed loop control) installation was designed over a twenty degree angle-of-incidence range. The Optimal Robust installation used economical Robust Design methodology to arrive at a single design which operated over the entire angle-of-incident range (open loop control). The Optimal Adaptive installation optimized all the design parameters at each angle-of-incidence. Thus, the Optimal Adaptive installation would require a closed loop control system to sense a proper signal for each effector and modify that effector device, whether mechanical or fluidic, for optimal inlet performance. In general, the performance differences between the Optimal Adaptive and Optimal Robust installation designs were found to be marginal. This suggests, however, that Optimal Robust open loop installation designs can be very competitive with Optimal Adaptive close loop designs. Secondary flow control in inlets is inherently robust, provided it is optimally designed. Therefore, the new methodology presented in this paper, combined array 'Lower Order' approach to Robust DOE, offers the aerodynamicist a very viable and economical way of exploring the concept of Robust inlet design, where the mission variables are brought directly into the inlet design process and insensitivity or robustness to the mission variables becomes a design objective.

Anderson, Bernhard H.↗
NASA NTRS2003

Task IV: Development of Circumferential Inlet Distortion through a Representative Eleven Stage High-speed Axial Compressor

The concepts and the procedure developed in Task I and Task III were used to determine the response of an eleven-stage high-speed compressor to an inlet distortion of 180 deg. circumferential extent for contrasting against its performance under uniform inlet flow. Using the computed results at the inlet to and outlet of the compressor, the computed total pressure ratio and efficiency for the clean condition are determined to be 14.22 and 76.9 percent respectively. As for the distorted case, these are determined to be 10.35 and 71.8 percent respectively, showing deterioration 76.9 percent vs 71.8 percent). The physical consistency of the computed flow field was assessed as a means of demonstrating the applicability and utility of the body force representation for inlet distortion computations. Specifically the computed evolution of the distorted pattern in static pressure and total pressure from compressor inlet to exit is examined. For the eleven-stage compressor examined here, the deterioration in performance has been found to be particularly severe in the last 2 stages. This suggests that the last two stages could be redesigned to alleviate the observed deterioration thus making the compressor performance insensitive to circumferential inlet distortion. This can potentially be accomplished by first determining what should the body force distribution of the last two stages should be to achieve minimal or no deterioration in performance in the last two stages. One can then in principle proceed to determine the blade design to yield such a body force distribution.

Tan, Choon-Sooi↗
NASA NTRS2003

The Role of Design-of-Experiments in Managing Flow in Compact Air Vehicle Inlets

It is the purpose of this study to demonstrate the viability and economy of Design-of-Experiments methodologies to arrive at microscale secondary flow control array designs that maintain optimal inlet performance over a wide range of the mission variables and to explore how these statistical methods provide a better understanding of the management of flow in compact air vehicle inlets. These statistical design concepts were used to investigate the robustness properties of low unit strength micro-effector arrays. Low unit strength micro-effectors are micro-vanes set at very low angles-of-incidence with very long chord lengths. They were designed to influence the near wall inlet flow over an extended streamwise distance, and their advantage lies in low total pressure loss and high effectiveness in managing engine face distortion. The term robustness is used in this paper in the same sense as it is used in the industrial problem solving community. It refers to minimizing the effects of the hard-to-control factors that influence the development of a product or process. In Robustness Engineering, the effects of the hard-to-control factors are often called noise , and the hard-to-control factors themselves are referred to as the environmental variables or sometimes as the Taguchi noise variables. Hence Robust Optimization refers to minimizing the effects of the environmental or noise variables on the development (design) of a product or process. In the management of flow in compact inlets, the environmental or noise variables can be identified with the mission variables. Therefore this paper formulates a statistical design methodology that minimizes the impact of variations in the mission variables on inlet performance and demonstrates that these statistical design concepts can lead to simpler inlet flow management systems.

Anderson, Bernhard H.↗
NASA NTRS2004

Effect of Blowing on Boundary Layer of Scarf Inlet

When aircraft operate in stationary or low speed conditions, airflow into the engine accelerates around the inlet lip and pockets of turbulence that cause noise and vibration can be ingested. This problem has been encountered with engines equipped with the scarf inlet, both in full scale and in model tests, where the noise produced during the static test makes it difficult to assess the noise reduction performance of the scarf inlet. NASA Langley researchers have implemented boundary layer control in an attempt to reduce the influence of the flow nonuniformity in a 12-in. diameter model of a high bypass fan engine mounted in an anechoic chamber. Static pressures and boundary layer profiles were measured in the inlet and far field acoustic measurements were made to assess the effectiveness of the blowing treatment. The blowing system was found to lack the authority to overcome the inlet distortions. Methods to improve the implementation of boundary layer control to reduce inlet distortion are discussed.

Gerhold, Carl H.↗
NASA NTRS1947

A Low-Speed Investigation of a Fuselage-Side Air Inlet for use at Transonic Flight Speeds

A low-speed investigation in the Langley propeller-research tunnel of annular air inlets designed to avoid compression shocks and attendant boundary-layer separation on the fuselage ahead of the inlets at transonic flight speeds by maintaining substream flow velocities on the fuselage nose was reported in NACA RM No. L6J04. In the present investigation, one of the original annular inlets was converted by the installation of a canopy and a nose-wheel fairing into a twin side inlet in order to study problems involved in applying such an inlet to a fighter-type airplane. Extensive measurements of pressures on the surface of the model and surveys of the internal flow were conducted at angles of attack of 0 degrees, 3 degrees, and 6 degrees over a wide range of inlet-velocity ratio.

Nichols, Mark R.↗
NASA NTRS2008

Boundary-Layer-Ingesting Inlet Flow Control

An experimental study was conducted to provide the first demonstration of an active flow control system for a flush-mounted inlet with significant boundary-layer-ingestion in transonic flow conditions. The effectiveness of the flow control in reducing the circumferential distortion at the engine fan-face location was assessed using a 2.5%-scale model of a boundary-layer-ingesting offset diffusing inlet. The inlet was flush mounted to the tunnel wall and ingested a large boundary layer with a boundary-layer-to-inlet height ratio of 35%. Different jet distribution patterns and jet mass flow rates were used in the inlet to control distortion. A vane configuration was also tested. Finally a hybrid vane/jet configuration was tested leveraging strengths of both types of devices. Measurements were made of the onset boundary layer, the duct surface static pressures, and the mass flow rates through the duct and the flow control actuators. The distortion and pressure recovery were measured at the aerodynamic interface plane. The data show that control jets and vanes reduce circumferential distortion to acceptable levels. The point-design vane configuration produced higher distortion levels at off-design settings. The hybrid vane/jet flow control configuration reduced the off-design distortion levels to acceptable ones and used less than 0.5% of the inlet mass flow to supply the jets.

Owens, Lewis R.↗
NASA NTRS2010

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Jet

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan has been completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 ft. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a three dimensional (3-D) code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

Chima, Rodrick V.↗
NASA NTRS2010

Computational Fluid Dynamics (CFD) Simulation of Hypersonic Turbine-Based Combined-Cycle (TBCC) Inlet Mode Transition

Methods of computational fluid dynamics were applied to simulate the aerodynamics within the turbine flowpath of a turbine-based combined-cycle propulsion system during inlet mode transition at Mach 4. Inlet mode transition involved the rotation of a splitter cowl to close the turbine flowpath to allow the full operation of a parallel dual-mode ramjet/scramjet flowpath. Steady-state simulations were performed at splitter cowl positions of 0deg, -2deg, -4deg, and -5.7deg, at which the turbine flowpath was closed half way. The simulations satisfied one objective of providing a greater understanding of the flow during inlet mode transition. Comparisons of the simulation results with wind-tunnel test data addressed another objective of assessing the applicability of the simulation methods for simulating inlet mode transition. The simulations showed that inlet mode transition could occur in a stable manner and that accurate modeling of the interactions among the shock waves, boundary layers, and porous bleed regions was critical for evaluating the inlet static and total pressures, bleed flow rates, and bleed plenum pressures. The simulations compared well with some of the wind-tunnel data, but uncertainties in both the windtunnel data and simulations prevented a formal evaluation of the accuracy of the simulation methods.

Slater, John W.↗
NASA NTRS2010

Boundary-layer-ingesting inlet flow control system

A system for reducing distortion at the aerodynamic interface plane of a boundary-layer-ingesting inlet using a combination of active and passive flow control devices is disclosed. Active flow control jets and vortex generating vanes are used in combination to reduce distortion across a range of inlet operating conditions. Together, the vortex generating vanes can reduce most of the inlet distortion and the active flow control jets can be used at a significantly reduced control jet mass flow rate to make sure the inlet distortion stays low as the inlet mass flow rate varies. Overall inlet distortion, measured and described as average SAE circumferential distortion descriptor, was maintained at a value of 0.02 or less. Advantageous arrangements and orientations of the active flow control jets and the vortex generating vanes were developed using computational fluid dynamics simulations and wind tunnel experimentations.

Owens, Lewis R.↗
NASA NTRS2009

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Aircraft

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan was completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 feet. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a 3-D code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

Chima, Rodrick V.↗
NASA NTRS2011

Large-Scale Low-Boom Inlet Test Overview

This presentation provides a high level overview of the Large-Scale Low-Boom Inlet Test and was presented at the Fundamental Aeronautics 2011 Technical Conference. In October 2010 a low-boom supersonic inlet concept with flow control was tested in the 8'x6' supersonic wind tunnel at NASA Glenn Research Center (GRC). The primary objectives of the test were to evaluate the inlet stability and operability of a large-scale low-boom supersonic inlet concept by acquiring performance and flowfield validation data, as well as evaluate simple, passive, bleedless inlet boundary layer control options. During this effort two models were tested: a dual stream inlet intended to model potential flight hardware and a single stream design to study a zero-degree external cowl angle and to permit surface flow visualization of the vortex generator flow control on the internal centerbody surface. The tests were conducted by a team of researchers from NASA GRC, Gulfstream Aerospace Corporation, University of Illinois at Urbana-Champaign, and the University of Virginia

Hirt, Stefanie↗
NASA NTRS2011

Inlet Flow Control and Prediction Technologies for Embedded Propulsion Systems

Fail-safe, hybrid, flow control (HFC) is a promising technology for meeting high-speed cruise efficiency, low-noise signature, and reduced fuel-burn goals for future, Hybrid-Wing-Body (HWB) aircraft with embedded engines. This report details the development of HFC technology that enables improved inlet performance in HWB vehicles with highly integrated inlets and embedded engines without adversely affecting vehicle performance. In addition, new test techniques for evaluating Boundary-Layer-Ingesting (BLI)-inlet flow-control technologies developed and demonstrated through this program are documented, including the ability to generate a BLI-like inlet-entrance flow in a direct-connect, wind-tunnel facility, as well as, the use of D-optimal, statistically designed experiments to optimize test efficiency and enable interpretation of results. Validated improvements in numerical analysis tools and methods accomplished through this program are also documented, including Reynolds-Averaged Navier-Stokes CFD simulations of steady-state flow physics for baseline, BLI-inlet diffuser flow, as well as, that created by flow-control devices. Finally, numerical methods were employed in a ground-breaking attempt to directly simulate dynamic distortion. The advances in inlet technologies and prediction tools will help to meet and exceed "N+2" project goals for future HWB aircraft.

McMillan, Michelle L.↗
NASA NTRS2012

Computational Analyses of the LIMX TBCC Inlet High-Speed Flowpath

Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for the high-speed flowpath and isolator of a dual-flowpath Turbine-Based Combined-Cycle (TBCC) inlet using the Wind-US code. The RANS simulations were performed in preparation for the Large-scale Inlet for Mode Transition (LIMX) model tests in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel. The LIMX inlet has a low-speed flowpath that is coupled to a turbine engine and a high-speed flowpath designed to be coupled to a Dual-Mode Scramjet (DMSJ) combustor. These RANS simulations were conducted at a simulated freestream Mach number of 4.0, which is the nominal Mach number for the planned wind tunnel testing with the LIMX model. For the simulation results presented in this paper, the back pressure, cowl angles, and freestream Mach number were each varied to assess the performance and robustness of the high-speed inlet and isolator. Under simulated wind tunnel conditions at maximum inlet mass flow rates, the high-speed flowpath pressure rise was found to be greater than a factor of four. Furthermore, at a simulated freestream Mach number of 4.0, the high-speed flowpath and isolator showed stability for freestream Mach number that drops 0.1 Mach below the design point. The RANS simulations indicate the yet-untested highspeed inlet and isolator flowpath should operate as designed. The RANS simulation results also provided important insight to researchers as they developed test plans for the LIMX experiment in GRC s 10- by 10-ft Supersonic Wind Tunnel.

Dippold, Vance F., III↗
NASA NTRS2012

Characterization of Aerodynamic Performance of Boundary-Layer-Ingesting Inlet Under Crosswind

NASA has been studying future transport concepts, envisioned to be technically realizable in the timeframe of 2020-2030, to meet environmental and performance goals. One concept receiving considerable interest involves a propulsion system embedded into a hybrid wing-body aircraft. While offering significant advantages in fuel savings and noise reduction by this concept, there are several technical challenges that are not encountered in the current fleet and must be overcome so as to deliver target performance and operability. One of these challenges is associated with an inlet system that ingests a significantly thick boundary layer, developing along the wing-body surface, into a serpentine diffuser before the flow meeting fan blades. The flow is subject to considerable total pressure loss and distorted at the fan face, much more significantly than in the inlet system of conventional aircraft. In our previous studies [1, 2], we have shown that through innovative design changes on the airframe surface, it is possible to simultaneously increase total pressure recovery and decrease distortion in the flow, without resorting to conventional penalty-ridden flow control concepts, such as vortex generator or boundary layer bleeding/suction. In the current study, we are interested in understanding the following issues: how the embedded propulsion system performs under a crosswind condition by studying in detail the flow characteristics of two inlets, the baseline and another optimized previously under the cruise condition. With the insight, it is hoped that it can help in the follow-on study by devising effective strategies to minimize flow distortion arising from the integration of an embedded-engine system into an airframe to the level acceptable to the operation and fuel consumption before 2030. To achieve these demanding goals, non-conventional concepts are called for; but technology gap is too big that it requires evolutionary approach by focusing various concepts and technologies needed in the next three generations of aircraft, respectively named as N+1, N+2, and N+3. Noticeably, considerable reduction in each category of 1 is required in N+2 (relative to Boeing 777-200 and GE90 engines) and N+3 (relative to Boeing 737-800 and CFM56-7B engines). In this study, concepts for N+2 is our interest. A concept that has potential to achieve these metrics and has been under intensive study is the hybrid wing body (HWB) airframe with a tightly integrated propulsion system, see Fig. 1. The inlet is non-circular at the entrance and the entering flow, no longer uniform or free of disturbances, and is now carrying with it a boundary layer developing along the fuselage; the inlet is thus known as boundary-layer-ingesting (BLI) inlet.

Liou, Meng-Sing↗
NASA NTRS2013

Combined Cycle Engine Large-Scale Inlet for Mode Transition Experiments: System Identification Rack Hardware Design

The System Identification (SysID) Rack is a real-time hardware-in-the-loop data acquisition (DAQ) and control instrument rack that was designed and built to support inlet testing in the NASA Glenn Research Center 10- by 10-Foot Supersonic Wind Tunnel. This instrument rack is used to support experiments on the Combined-Cycle Engine Large-Scale Inlet for Mode Transition Experiment (CCE‒ LIMX). The CCE‒LIMX is a testbed for an integrated dual flow-path inlet configuration with the two flow paths in an over-and-under arrangement such that the high-speed flow path is located below the lowspeed flow path. The CCE‒LIMX includes multiple actuators that are designed to redirect airflow from one flow path to the other; this action is referred to as "inlet mode transition." Multiple phases of experiments have been planned to support research that investigates inlet mode transition: inlet characterization (Phase-1) and system identification (Phase-2). The SysID Rack hardware design met the following requirements to support Phase-1 and Phase-2 experiments: safely and effectively move multiple actuators individually or synchronously; sample and save effector control and position sensor feedback signals; automate control of actuator positioning based on a mode transition schedule; sample and save pressure sensor signals; and perform DAQ and control processes operating at 2.5 KHz. This document describes the hardware components used to build the SysID Rack including their function, specifications, and system interface. Furthermore, provided in this document are a SysID Rack effectors signal list (signal flow); system identification experiment setup; illustrations indicating a typical SysID Rack experiment; and a SysID Rack performance overview for Phase-1 and Phase-2 experiments. The SysID Rack described in this document was a useful tool to meet the project objectives.

Thomas, Randy↗
NASA NTRS2015

Method for Determining Optimum Injector Inlet Geometry

A method for determining the optimum inlet geometry of a liquid rocket engine swirl injector includes obtaining a throttleable level phase value, volume flow rate, chamber pressure, liquid propellant density, inlet injector pressure, desired target spray angle and desired target optimum delta pressure value between an inlet and a chamber for a plurality of engine stages. The method calculates the tangential inlet area for each throttleable stage. The method also uses correlation between the tangential inlet areas and delta pressure values to calculate the spring displacement and variable inlet geometry of a liquid rocket engine swirl injector.

Trinh, Huu P.↗
NASA NTRSDate not supplied

Inlet Vortex Generator Design for the X-59 Low Boom Flight Demonstrator

An array of vortex generators was designed for the X-59 Low Boom Flight Demonstrator aircraft inlet. Prior to the a High Speed Inlet Distortion Test, concern regarding potentially high measurements of steady state or dynamic distortion prompted an effort to examine the use of passive flow control to reduce distortion levels. Design of experiments methodology was used to develop two matrices of vortex generators placed inside the inlet serpentine duct and upstream of the inlet on the compression surface, varying on several geometric parameters. Reynolds-Averaged Navier-Stokes simulations were performed using the FUN3D flow solver at Mach 1.45 to analyze the impact of each configuration on total pressure recovery, circumferential distortion, and radial distortion. Results from simulations of the baseline aircraft are compared with data acquired in the NASA Glenn Research Center’s 8- by 6-foot Supersonic Wind Tunnel. A vortex generator configuration with vanes placed upstream of the inlet terminal shock was shown to significantly increase total pressure recovery and decrease circumferential distortion with no penalty to radial distortion levels. Inlet mass flow sweeps at Mach 1.45, 0.6, and 0.3 indicated no detrimental effects from the vanes at off-design conditions.

Vortex generators↗