Aircraft Engine Inlet Research Articles

Investigation of the Internal Flow Characteristics of a Tiltrotor Aircraft Engine Inlet in a Gust Environment

In the vertical take-off and landing (VTOL) state of tiltrotor aircraft, the inlet entrance encounters the incoming airflow at a 90° attack angle, resulting in highly complex internal flow characteristics that are extremely susceptible to gusts. Meanwhile, the flow quality at the inlet exit directly affects the performance of the aircraft’s engine. This work made use of an unsteady numerical simulation method based on sliding meshes to investigate the internal flow characteristics of the inlet during the hover state of a typical tiltrotor aircraft and the effects of head-on gusts on the inlet’s aerodynamic characteristics. The results show that during the hover state, the tiltrotor aircraft inlet features three pairs of transverse vortices and one streamwise vortex at the aerodynamic interface plane (AIP). The transverse vortices generated due to the rotational motion of the air have the largest scale and exert the strongest influence on the inlet’s performance, which is characterized by pronounced unsteady features. Additionally, strong unsteady characteristics are present within the inlet. Head-on gusts mainly affect the mechanical energy and non-uniformity of the air sucked into the inlet by influencing the direction of the rotor’s induced slipstream, thereby impacting the performance of the inlet. The larger head-on gusts have beneficial effects on the performance of the inlet. When the gust velocity reaches 12 m/s, there is a 1.01% increase in the total pressure recovery (σ) of the inlet, a 25.72% decrease in the circumferential distortion index (DC60), and a reduction of 62.84% in the area where the swirl angle |α| exceeds 15°. Conversely, when the gust velocity of head-on gusts reaches 12 m/s in the opposite direction, the inlet’s total pressure recovery decreases by 1.13%, the circumferential distortion index increases by 14.57%, and the area where the swirl angle exceeds 15° increases by 69.59%, adversely affecting the performance of the inlet. Additionally, the presence of gusts alters the unsteady characteristics within the inlet.

Modeling and Analysis of Anti-Icing Power of Aircraft Engine Inlet Based on Symmetric Algorithms

To evaluate the capability of engine inlet, inlet components and power plant anti ICER under low temperature, this paper introduces the evaluation method of anti icing system for civil aviation engine room, and analyzes the anti icing power of the aircraft intake based on the symmetric algorithm. The realizable k-cube model and wall function method are used to analyze the flow field in the inlet of an aircraft engine. Based on the analysis of the flow field of the intake port of an aircraft engine, the anti ice power of the intake port is calculated according to the heat balance relationship of the intake port surface. The symmetrical particle swarm algorithm is adopted to optimize the calculation process of inlet anti-ice power, and the particle wide area learning strategy is used to promote the calculation of inlet anti-ice power. In this way, the computational complexity is significantly reduce and the accuracy of the power analysis of the inlet anti-ice is enhanced. The simulation results show that the absolute error of the proposed method is less than 1% in 1000 iterations. Through the analysis of the surface temperature changes of the inlet deflector under different experimental conditions, it can be known that the method can effectively analyze the anti-icing power of aircraft engine inlet.

Characterization of gas turbine ingested sand particles based on electrostatic signal

Particle ingestion into a gas turbine can have serious effects on both performance and engine in-service reliability. Thus there exists a need for in situ monitoring and characterizing particulate matter entering an aircraft engine inlet for the purposes of engine damages estimation and prognosis. This paper presents the initial development of Ingested Debris Monitoring System (IDMS) signal processing method of characterizing the ingested particles. A theoretical analysis and simulation study were carried out to study the relationships between the characteristics of the ingested sand particles and the features of the IDMS signal both in frequency- and time-domain. A Finite-Element Modeling (FEM) for the IDMS Sensor was developed, then the validated FEM modeling was used for simulation experiments of particles ingestion under various conditions of different particle moving speeds, concentrations and charge-to-mass ratios. Results of the theoretical analysis and simulation study demonstrates the feasibility and effectiveness of the proposed method to provide real time information characterizing the size and concentration of ingested sand particles, and will serve as an impetus to carry out further research.

CFD analysis of conical diffuser under swirl flow inlet conditions using turbulence models

Diffuser are used to maximize static pressure recovery in flow systems such as turbo-machines, turbine exit, duct flows, aircraft engine inlets, closed circuit wind tunnels etc. Many of turbo-machines has swirl flow at inlet which requires thorough investigation to predict fluid flow parameters like velocity, pressure and mass flow. The current research investigates the conical diffuser subjected to swirl flow inlet conditions using Computational Fluid Dynamics. Different cone angles of diffuser are considered for analysis i.e. 10°, 20° and 30°. The CAD model of diffuser is developed using Creo 2.0 software and CFD analysis is conducted using RNG k-epsilon turbulence models using ANSYS CFX software. The results generated from CFD analysis has thrown light on effect of diffuser cone angles with respect to outlet pressure, velocity and turbulence kinetic energy. The research shows that cone angle has significant effect on static pressure recovery of diffusers.

The experimental investigation of water spray disperse parameters behind the water injecting system at the aircraft engine inlet by applying PSV

To provide an aircraft engine certification tests for water ingestion, it is necessary to control accurately the parameters of the water drops generated by the water injecting device simulating rain at the engine inlet. The paper presents the results of autonomous tests of a single-channel centrifugal nozzle for the water injecting device. The tests purpose is to investigate the disperse parameters of the water spray for the typical operation modes of the nozzle as a part of the water injecting system. Tests were carried out at the experimental setup using the PSV (Particle Shadow Velocimetry) technique at three typical nozzle operation modes. As a result of the tests, the distribution graphs of the D10, D32 and Dv50 diameters, axial velocity (U) and mass flow (Mabs) of water droplets over the radius of the spray pattern have been obtained. To assess Dv50 mean median diameter of spray droplets the cumulative distribution functions of the droplets mass fraction depending on the size for each of the operation modes of the nozzle have been plotted. On the basis of these functions, the mean median diameters of water spray droplets have been calculated.

THE INFLUENCE OF THE ANGLE OF ATTACK ALPHA ON THE PRESSURE RECOVERY AT AERODYNAMICS INTERFACE PLANE INVESTIGATION

In the process of design, anaircraft to know the performance and stability requires testing in a wind tunnel.Similarly, the selection of engines for aircraft design embedded on conducted tests in a wind tunnel model air intake with some variation of the angle of attack.This paper examines the effect ofthe high angle of attack alpha testing aircraft air intake with a position of beta 0° on the subsonic regime with wind speeds of 65 ms-1. Combat aircraft flying with an unexpected manoeuvre to avoidenemy attack nor did defence against missiles.Often the aircraft must fly to the position of a very high angle of attack.The theoretical value of the Pressure Recovery is approaching one in order to make the machine pinned can work optimally.Experimental observations of Pressure Recovery plotted against the variation of the angle of attack alpha.The observations show the performance design of aircraft engine inlet on the angle of attack alpha large can still give adequate pressure recovery.

A numerical investigation of Helmholtz resonators in the presence of grazing flow by means of the lattice Boltzmann method

Helmholtz resonators remain widely used in noise control. In applications such as aircraft engines and exhaust systems, the presence of a grazing flow significantly changes their behavior, and a correct prediction of their acoustic properties is essential to improve noise reduction. With the purpose of understanding the physical phenomena associated with the acoustic-flow interaction, the simulation of a single 2D Helmholtz resonator was considered by means of an in-house numerical code based on the lattice Boltzmann method (LBM). The results have been validated against published data based on both experimental results and direct numerical simulation (DNS) for normally incident acoustic waves in the absence of flow. The investigations will proceed by taking into account grazing acoustic waves in the presence of a grazing flow, similarly to the conditions found in a liner test rig. Efforts will be focused on typical aircraft engine inlet conditions, i.e., high Mach numbers and SPL. Both experimental and numerical results will be compared in terms of absorption coefficient and impedance.

Flow Control by Pulsed Jet in a Curved -Duct: A Spectral Analysis

This paper presents an experiment of flow control by supersonic pulsed jets in a very aggressive -duct typical of modern fighter aircraft engine inlet ducts. The Mach number in the aerodynamic interface plane ranges from 0.2 to 0.4, which is realistic from an application point of view. It was shown that, with continuous blowing, the flow could be fully reattached. The sensitivity of the distortion coefficient to the jet frequency was also clearly evidenced. Surprisingly, it is recommended to avoid actuating the jets at the natural frequencies present in the separated zone of the uncontrolled flow. Besides, it is observed that pulsed jets actuation increases the unsteady flow distortion. Furthermore, in the uncontrolled flow, a lateral movement of the structures associated with high total pressure losses in the aerodynamic interface plane was highlighted.

Investigation of foreign object throw by reversal jets to the aircraft engine inlet

A model of reversal jet interaction with a solid particle and particle throw from the runway surface to the GTE inlet is proposed. The calculation results are compared with experimental data on foreign object throw from the runway surface in the Tu-154 airplane landing. We present the regularities that make it possible to develop design and operational measures to improve protection of engines with thrust reversal when they are located behind the wing at the fuselage sides.

Active Control of Low Frequency Buzz-Saw Tones; Theory and Experiment

This paper discusses some theoretical and experimental issues related to the active control of buzz-saw tones radiated from an aircraft engine inlet at takeoff. Buzz-saw tones occur when the rotor tip speed becomes supersonic. It arises due to the formation of shocks at the blade tips, which then propagate non-linearly along the engine duct. Buzz-saw noise comprises a series of tones harmonically related to the shaft rotation frequency (Engine Orders). Moreover, it has been shown that the noise at each Engine Order comprises just one single spinning mode. Whilst there has been considerable work aimed at assessing the feasibility of active control for reducing the tones due to rotor – stator interaction, buzz-saw noise has been overlooked. This is surprising given the low frequencies at which Buzz-Saw noise occurs, and its simple modal content, which both suggest that buzz-saw noise is particularly suited to active control. This paper describes various control strategies for reducing buzz-saw noise. The paper concludes with the results of a laboratory experiment aimed at demonstrated experimentally the effectiveness with which a single spinning mode, characteristic of an EO = 3 buzz-saw tone, can be controlled. Reductions in excess of 14dB are reported at some frequencies.

Preconditioning Harmonic Unsteady Potential Flow Calculations

This paper considers finite element discretizations of the Helmholtz equation and its generalization arising from harmonic acoustic perturbations to a nonuniform steady potential flow. A novel elliptic, positive-definite preconditioner with a multigrid implementation is used to accelerate the iterative convergence of Krylov subspace solvers. Both theory and numerical results show that for a model 1-D Helmholtz test problem, the preconditioner clusters the discrete system's eigenvalues and lowers its condition number to a level independent of grid resolution. For the 2-D Helmholtz equation, grid-independent convergence is achieved using a quasi-minimal residual Krylov solver, significantly outperforming the popular symmetric successive over-relaxation preconditioner. Impressive results are also presented on more complex domains, including an axisymmetric aircraft engine inlet with nonstagnant mean flow and modal boundary conditions.

Radar Cross-Section Measurements of a Full-Scale Aircraft Duct/Engine Structure

Cavity radar cross-section (RCS) data are measured from a full-sized generic aircraft duct/engine mock-up experimental apparatus. Details of the experimental measurements and characterization of the RCS from a generic aircraft engine inlet with a fan assembly embedded inside are described. The experimental results reveal that the azimuthal RCS patterns from an aircraft engine inlet with fan components embedded inside are very complex

Numerical Study of Approximate Inverse Preconditioner for Two-Dimensional Engine Inlet Problems

The method of moments has been widely used in the analysis of electromagnetic scattering problems. Recently, significant advances in the development of fast and efficient techniques for solving large problems have been reported. In such methods, iterative matrix solvers are preferred by virtue of their speed and low memory requirements. But for near resonant and strong multiple scattering problems, e.g., involving an aircraft engine inlet, a large number of iterations is required for convergence. In this paper, an efficient approximate inverse based preconditioner is used to reducing this number of iterations. By using the matrix partitioning method, the computational cost for obtaining the approximate inverse is reduced to O(N). We apply this preconditioner to an O(N log N) algorithm, the multilevel fast multipole algorithm, for the aircraft engine inlet problem. The numerical results show the efficiency of this preconditioner.

Active control of aircraft engine inlet noise using compact sound sources and distributed error sensors

Active control of aircraft engine inlet noise using compact sound sources and distributed error sensors

Active control of aircraft engine inlet noise using compact sound sources and distributed error sensors

An active noise control system using a compact sound source is effective to reduce aircraft engine duct noise. The fan noise from a turbofan engine is controlled using an adaptive filtered-x algorithm. Single, multi channel control systems are used to control the fan blade passage frequency (BPF) tone and the BPF tone and the first harmonic of the BPF tone for a plane wave excitation. The multi channel control system is used to control fan tones and a high pressure compressor BPF tone simultaneously, and any spinning mode. A compact sound source is employed to generate the control field. This compact sound source consists of an array of identical thin, cylindrically curved panels (125) with an inner radius of curvature corresponding to that of the engine inlet. These panels are flush mounted inside the inlet duct (Inlet Wall) and sealed on all edges to prevent leakage around the panel.

Axially Symmetric Acoustic Wave Propagation Through Flows with Vorticity

Periodic solutions of the linearized axially symmetric Euler equations are sought. The axially symmetric case is of considerable practical interest, since these equations are used to calculate the sound field around commercial aircraft engine inlets, which are circular. It is known that these equations are well posed in plane case; however, the proof relies on the two-dimensional nature of the problem. The axially symmetric problem does not trivially follow, since periodic variations in the $\theta $ direction give rise to a three-dimensional velocity field.

Water droplet impingement on airfoils and aircraft engine inlets foricing analysis

This paper includes the results of a significant research program for verification of computer trajectory codes used in aircraft icing analysis. Experimental water droplet impingement data have been obtained in the NASA Lewis Research Center Icing Research Tunnel for a wide range of aircraft geometries and test conditions. The body whose impingement characteristics are required is covered at strategic locations by thin strips of moisture absorbing (blotter) paper and then exposed to an airstream containing a dyed-water spray cloud. Water droplet impingement data are extracted from the dyed blotter strips by measuring the optical reflectance of the dye deposit on the strips with an automated reflectometer. Impingement characteristics for all test geometries have also been calculated using two recently developed trajectory computer codes. Good agreement is obtained with experimental data. The experimental and analytical data show that maximum impingement efficiency and impingement limits increase with mean volumetric diameter for all geometries tested. For all inlet geometries tested, as the inlet mass flow is reduced, the maximum impingement efficiency is reduced and the location of the maximum impingement shifts toward the inlet inner cowl.

Collocation for an integral equation arising in duct acoustics

A mathematical model is developed to describe the effect of aircraft-engine inlet geometry on the reflected and radiated acoustic field without flow, as studied experimentally using a spinning-mode synthesizer by Silcox (1983). The acoustic pressure in the inlet interior and exterior is modeled by a pure cylindrical azimuthal mode for the Helmholtz equation with hardwall boundary and by the Helmholtz equation and the radiation condition at infinity, respectively. The analytical approach to the solution of the resulting boundary-value problem and the program implementation are explained; numerical results are presented in tables and graphs; and the uniqueness of the problem is demonstrated.

Spinning mode sound propagation in ducts with acoustic treatment

Recent acoustic data have shown large noise attenuations for acoustically treated aircraft engine inlets without splitter rings. These encouraging data have stimulated a more detailed theoretical study of the acoustic propagation of spinning modes in acoustically treated open circular ducts. In addition, the suppressor with splitter rings was modeled by using the rectangular approximation to the annular duct. The theoretical models were used to optimize the suppressor designs for several spinning lobe numbers from zero to 50. Some interesting results of the analysis are that for circular ducts the maximum possible attenuation and the optimum wall impedance are strong functions of the lobe number. For annular ducts, the attenuation and optimum wall impedance are insensitive to the spinning lobe number for well-cut-on modes. The above results help explain why suppressors with splitter rings have been quite effective in spite of the lack of detailed information on the noise-source modal structure. Conversely, effective use of wall treatment only will require expanded knowledge of the noise-source structure. Approximate solutions are presented to help in understanding the more exact theoretical results.