Engine Fan Research Articles

Initial data for the calculation of fluid flywheel for internal combustion engine fan drive

Procedure of initial data generation for calculation and designing of fluid flywheel for internal combustion engine fan drive is given. Fans can be devices of air or fluid cooling, and cooling draught inlet can be axial and radial-axial. The central purpose is to generate algorithm and software for personal computer with maximal exploitation of analytical dependences instead of graphs.

Design and Analysis of Specimen and Device for Biaxial CCF Experiment

Experiment devices and specimens are designed to simulate the combined cycle fatigue (CCF) phenomena occurred in fan and compressor blades of gas turbine engine. A uniaxial fatigue test machine is modified to conduct biaxial loadings through adding an electromagnetic shaker. Two plate specimens with or without middle platform are designed for CCF testing. The middle point of the specimen is chosen as the exciting point and two pins are used so that the interference between the two loading directions will be released. By means of FEA, it is approved that the system is suitable for CCF test of Titanium alloy and the similar materials.

Tribology in Turbo Fan Engines for Small Business Jets

Tribology in Turbo Fan Engines for Small Business Jets

Modeling the Super Plastic Forming of a Multi-Sheet Diffusion Bonded Titanium Alloy Demonstrator Fan Blade

The paper describes a finite element method in 2D and 3D to simulate the super plastic forming of a demonstrator jet engine fan blade made from Titanium alloy sheet. The fan blade is an assembly of three sheets in which a single inner (core) sheet is diffusion bonded to the two outer (skin) sheets at prescribed zones, which is then super-plastically formed to a desired fan profile. In the model, the diffusion bonded zones between the core and skin sheets are simulated using tied interfaces. The thickness of each skin sheet is not uniform and significant change in thickness can occur over a short distance as well as gradually over the entire skin sheet. The thickness of the core sheet which is smaller than the thickness of each skin sheet remains uniform. The paper describes the design for a scaled-down demonstrator fan blade and model build process. Selected results and evaluations of finite element simulations are presented and discussed.

Bladed disc crack diagnostics using blade passage signals

One of the major potential faults in a turbo fan engine is the crack initiation and propagation in bladed discs under cyclic loads that could result in the breakdown of the engines if not detected at an early stage. Reliable fault detection techniques are therefore in demand to reduce maintenance cost and prevent catastrophic failures. Although a number of approaches have been reported in the literature, it remains very challenging to develop a reliable technique to accurately estimate the health condition of a rotating bladed disc. Correspondingly, this paper presents a novel technique for bladed disc crack detection through two sequential signal processing stages: (1) signal preprocessing that aims to eliminate the noises in the blade passage signals; (2) signal postprocessing that intends to identify the crack location. In the first stage, physics-based modeling and interpretation are established to help characterize the noises. The crack initiation can be determined based on the calculated health monitoring index derived from the sinusoidal effects. In the second stage, the crack is located through advanced detrended fluctuation analysis of the preprocessed data. The proposed technique is validated using a set of spin rig test data (i.e. tip clearance and time of arrival) that was acquired during a test conducted on a bladed military engine fan disc. The test results have demonstrated that the developed technique is an effective approach for identifying and locating the incipient crack that occurs at the root of a bladed disc.

Psychoacoustic assessment of a new aircraft engine fan noise synthesis method

Simulation of aircraft flyover events can facilitate psychoacoustic studies exploring the effects of noise generated by future aircraft designs. The perceived realism of a simulated flyover event may be impacted by the perceived realism of the synthesized fan noise of the aircraft engine. Short-term fluctuations in tonal amplitude and frequency are important cues contributing to that perception of realism, but are not accounted for by predictions based on long-term averages. A new synthesis method has been developed at NASA Langley Research Center to generate realistic aircraft engine fan noise using predicted source noise directivities in combination with short-term fluctuations. In the new method, fluctuations in amplitude and frequency are included based upon analysis of static engine test data. Through psychoacoustic testing, this study assessed perceived effectiveness of the new synthesis method in generating realistic fan noise source. Realism was indirectly assessed by judging the similarity of synthesized sounds (with and without fluctuations) with recordings of fan noise. Results of ANOVA analyses indicated that subjects judged synthesized fan noise with fluctuations as being more similar to recordings than synthesized fan noise without fluctuations.

Ballistic Impact Testing of Aluminum 2024 and Titanium 6Al-4V for Material Model Development

An experimental program is underway to develop a consistent set of material property and impact test data, and failure analysis, for a variety of materials that can be used to develop improved impact failure and deformation models. Unique features of this set of data are that all material property information and impact test results are obtained using identical materials, the test methods and procedures are extensively documented and all of the raw data is available. This report describes ballistic impact testing which has been conducted on aluminum (Al) 2024 and titanium (Ti) 6Al-4vanadium (V) sheet and plate samples of different thicknesses and with different types of projectiles, one a regular cylinder and one with a more complex geometry incorporating features representative of a jet engine fan blade.

Psychoacoustic study on contribution of fan noise to engine noise

There are more researches on engine fan noise control focusing on reducing fan noise level through optimizing fan structure, and a lot of research achievements have been obtained. However, researches on the effect of fan noise to engine noise quality are lacking. The influences of the effects of fan structure optimization on the engine noise quality are unclear. Thus, there will be a decline in fan noise level, but the deterioration of engine noise quality. Aiming at the above problems, in consideration of fan structure design and engine noise quality, an innovative method to analyze the contribution of fan noise to engine noise quality using psychoacoustic theory is proposed. The noises of diesel engine installing different cooling fans are measured by using the acoustic pressure method. The experiment results are regarded as analysis samples. The model of sensory pleasantness is used to analyze the sound quality of a diesel engine with different cooling fans. Results show that after installing 10-blade fan in medium diameter the sensory pleasantness at each test point is increased, and the increase is 13.53% on average, which indicate the improvement of the engine noise quality. In order to verify the psychoacoustical analysis, the subjective assessment is carried out. The test result shows the noise quality of engine installed 10-blade fan in medium diameter is most superior. 1/3 octave frequency spectrum analysis is used to study the reason of the improvement of engine noise quality. It is found that after installing proper cooling fan the sound pressure level below 400 Hz are obviously increased, the frequency assignment and spectral envelope are more reasonable and a proper cooling fan can optimize the spectrum structure of the engine noise. The psychoacoustic study is applied in the contribution of fan noise to engine noise, and the idea of engine sound quality improvement through the structure optimization is proposed.

Multiple Sliding Modes with Override Logic: Limit Management in Aircraft Engine Controls

G AS turbine engines are subjected to several operational constraints, both at steady state and in the transient regime. These constraints correspond to mechanical limits such as speed and acceleration; limits in thermodynamic variables such as pressures and temperatures; and limits preventing engine components from undergoing dynamic instabilities, mainly surge, stall and flame blowout. Engine control systems incorporate measures to protect the engine from crossing these limits at all times. Achieving limit protection while ensuring fast transient response represents a design challenge, which was summarized by Spang and Brown in their 1999 paper [1]: “. . .much of the complexity of the control comes from the need to operate the engine as close as possible to its limits.” This paper presents a novel and very effective way to achieve such maximal exploitation of the available limits to enable fast responses in thrust or other relevant variables. The proposed technique is based on thewidely usedmultiregulator architecture with max–min limit protection logic [1,2], which uses linear compensators in its original form. As shown in Fig. 1, this architecture uses a single control input (fuel flow) to drive the desired output (typically fan speed or engine pressure ratio) to a setpoint through a control loop with integral action and a linear compensator. Additional compensators are set up to control limited variables to their permissible values, still using fuel flow as the control input. A max–min selection logic stage decides whether the main output regulation task or any of the limit protection tasks should be made active at anygiven time. Integral action is used to achieve type 1 loops for offset-free regulation. The regulators, thus, provide control input rates. Define two index sets as L f1; 2; . . . lg and H fl 1; l 2; . . . hg, where L contains the indices of the regulators associated with the min selector, while the elements of H are the indices of the regulators associatedwith themax selector. No index is defined for the output of the min selector. As shown in Fig. 1, the control rate applied to the integrator is selected according to the formula

The parametered modeling technology study for the civil aircraft engine fan blade

PurposeThe purpose of this paper is to introduce parametered modeling technology for the civil aircraft engine fan blade, to design the fan blade rapidly and accurately.Design/methodology/approachThe entire fan blade consists of three crucial parts: blade airfoil, tenon and airfoil root. Blade airfoil with a free surface feature is formed through the blade profiles from the hub to tip in the radial direction. The non‐uniform rational basis spline (NURBS) is utilized to describe the blade profile. The geometry model of fan blade tenon is generated by extruding the sketch of the tenon. And the fillet section is designed to achieve the smooth transition of the up surface and the bottom surface of the blade root. Furthermore, the fan blade of a typical commercial engine is redesigned by the above method.FindingsThe stress analysis of the fan blade shows that the fan blade model designed in this work is reasonable.Originality/valueThe parametered fan blade model is presented on the basics of feature‐based modeling technology.

A method to assess the effects of engineering change propagation

This paper presents a modelling method that seeks to support the prediction and management of undesired engineering change (EC) propagation during the design and development of complex products. The method builds on the house of quality and the change prediction method to model the effects of potential change propagation brought about by different change options. The objective is to better reflect how well each change option can address the product requirements. The method was applied to the design of a jet engine fan during a case study with an aerospace company. The findings suggest that this modelling approach is suitable for assessing the effects of potential EC propagation and can support companies in effective exploration of the design space.

Hybrid contact stress analysis of a turbine engine blade to disk attachment

The present investigation examines an analysis methodology for fretting fatigue in a turbine engine fan disk. This is an important problem for the operators of turbine engines, since it is a significant driver of fatigue damage and failure risk of disks. Fretting fatigue in turbine engines occurs when the blade and disk are pressed together in contact and experience a small oscillating relative displacement due to variations in engine speed and vibratory loading. Fretting causes a very high local stress near the edge of contact resulting in wear, nucleation of cracks, and their growth, which can result in significant reduction in the life of the material. It is dependent on geometry, loading conditions, residual stresses, and surface roughness, among other factors. These complexities are not just physically based, but also computationally challenging. For example, the determination of the local contact stresses accurately depends on the mesh resolution of the finite element method (FEM) model. This has been addressed using various approaches. Recently, a computational hybrid technique was implemented successfully to investigate fretting fatigue of turbine engine blade and disk attachments. The present work extends application to specifically investigate the effects of surface contact in an actual blade and disk assembly using a representative loading mission. The results show consistency with available experimental data. Finally, the knowledge gained from this investigation could be used as a basis for uncertainty analyses of an actual blade and disk assembly.

A Statistical Method for Crack Detection from Vibrothermography Inspection Data

Nondestructive evaluation methods are widely used in quality control processes for critical components in systems such as aircraft engines and nuclear power plants. These same methods are also used in periodic field inspection to assure that the system continues to be reliable while it is in service. This paper describes a detection algorithm to automatically analyze vibrothermography sequence-of-image inspection data used to detect cracks in jet engine fan blades. Principal components analysis is used for dimension reduction. Then the fitted coefficients of the first principal component are processed by using robust regression to produce studentized residuals that are used in the crack-detection rules. We also show how to quantify the probability of detection for the algorithm. The detection algorithm is both computationally efficient and accurate. It correctly identified several cracks in our experimental data that were not detected by the standard human visual inspection method.

SPH Method Based on Bird Impact Numerical Simulation of Aircraft Engine Fan

The traditional numerical methods about calculation of bird strike problems are Lagrangian numerical method and Euler method. However, the existence of the finite element mesh distortion and the problem of insufficient accuracy appeared respectively. In this paper, the SPH meshless method applied to simulate bird strike aircraft engine fan, by comparing the simulated data and experimental data to test the accuracy of the simulation about bird hit the fan that is used SPH method .And by comparing the SPH method and traditional methods of bird strike problems in the time step to prove the superiority of the SPH method.

Rotor Dynamic Balance on the Field for Cooling Fan of Engine

The paper analyzes the rotor imbalance of cooling fan for the engine, and come to influence coefficient per unit quality. We can easy count the practice balancer cooling fan rotor. This method is very simple.

Development of a transonic wind tunnel to investigate engine bypass flow heat exchangers

The design of a new transonic wind tunnel to reproduce bypass aero-engine flow conditions is described. This study has been driven by the interest to investigate and optimize the aero-thermal effects of an air/oil integrated surface heat exchanger placed at the inner wall of the secondary duct of a turbofan. The test section of the intermittent wind tunnel consists of a complex three-dimensional (3D) geometry. The flow evolves over the splitter, duplicated at real scale, as it does in the engine, guided by helicoidally shaped lateral walls. Numerical flow analyses have been used to guide the design of the complete wind tunnel. Zero-dimensional models have been developed to recreate the wind tunnel behaviour. Two- and three-dimensional computations have been performed to optimize the test section and the inlet guide vanes that duplicate the flow downstream of the engine fan. The reproduction of the bypass flow structure has been numerically analysed. Flow computations with and without the presence of instrumentation in the test section have been contrasted. Additionally, the influence of an aerodynamic probe on the flow has been studied numerically. This study should lead to improvements in the thermal management of future aircraft power plants, particularly by taking benefit of the cold bypass air to refrigerate the lubrication oil, without penalizing the propulsive efficiency.

THREE-DIMENSIONAL FLOW IN A TRANSONIC AXIAL FLOW FAN OF A HIGH BYPASS RATIO TURBOFAN ENGINE

The present study deals with the air flow through an axial flow transonic fan blades, found in high bypass turbofan engines. Examples are CF-6, RB211 and GE90 turbofan engines. The fan has a wide chord, highly twisted blades of large span with a mean hub-to-tip ratio of about 0.404. The blades are tapered with its maximum chord at the tip radius. Moreover, these twisted blades have slight reduction in its maximum thickness from hub to tip. Consequently, the flow field is a three dimensional viscous compressible. The engine intake together with the fan resemble the case studied here. The computational domain includes two periodic merged sectors at a central angle of 360 degrees divided by the number of blades (38). The first sector is stationary, which represents the intake domain, while the second is rotating, which represents the fan domain that comprehending one blade. This is an effective way of reducing the calculation time and speeding up the iterations. The investigated model (Intake and fan blade) dimensions and configuration are identified by measuring real dimensions of CF-6 engine fan module. The FLUENT solver employing the Spalart-Allmaras turbulence model is used to solve the flow field in both the intake and fan. The program results represent the variation of the flow characteristics (pressure, temperature, Mach number, etc) in the flow field at the design condition (cruise conditions; namely, 11000 m altitude, 0.85 Mach number and 3674 rpm fan speed). The fan performance map was also predicted for a range of percentage fan corrected speed from 60% to 110%.

A computationally efficient non-linear beam model

Aerospace structures with large aspect ratio, such as airplane wings, rotorcraft blades, wind turbine blades, and jet engine fan and compressor blades, are particularly susceptible to aeroelastic phenomena. Finite element analysis provides an effective and generalized method to model these structures; however, it is computationally expensive. Fortunately, the large aspect ratio of these structures is exploitable as these potential aeroelastically unstable structures can be modeled as cantilevered beams, drastically reducing computational time. In this paper, the non-linear equations of motion are derived for an inextensional, non-uniform cantilevered beam with a straight elastic axis. Along the elastic axis, the cross-sectional center of mass can be offset in both dimensions, and the principal bending and centroidal axes can each be rotated uniquely. The Galerkin method is used, permitting arbitrary and abrupt variations along the length that require no knowledge of the spatial derivatives of the beam properties. Additionally, these equations consistently retain all third-order non-linearities that account for flexural–flexural–torsional coupling and extend the validity of the equations for large deformations. Furthermore, linearly independent shape functions are substituted into these equations, providing an efficient method to determine the natural frequencies and mode shapes of the beam and to solve for time-varying deformation. This method is validated using finite element analysis and is extended to swept wings. Finally, the importance of retaining cubic terms, in addition to quadratic terms, for non-linear analysis is demonstrated for several examples.

Finite Element Analysis on the Impact-induced Damage of Composite Fan Blades Subjected to a Bird Strike

Carbon fiber-reinforced composites have been recently applied for engine fan blades because of their high specific strength. In the design of the fan blade, bird-strike impact is one of the greatest concerns, since impact-induced damage can lead to the engine stalling. This study presents a numerical method to analyze bird-strike impact as a soft-body impact on a cantilevered composite panel. Especially, we coupled a stabilized dynamic contact analysis, which enables appropriate prediction of impact force on the panel, with laminate damage analysis to predict the impact-induced progressive damage in the composite. This method is verified through a comparison with experimental results. With the numerical method, we investigate the effect of impact condition, blade thickness and shape on the impact-induced damage in a composite fan blade subjected to a bird strike. An intermediate blade thickness and a large blade curvature help to improve the bird-striking impact resistance of the composite.

Measurements and Simulations of Wave Propagation for Wireless Sensor Networks in Jet Engine Turbines

In this letter, measurements and simulations of wave propagation inside a jet engine fan have been performed. The investigation was done using both EM simulations of different cases of propagation inside the engine and by measuring the corresponding cases inside a half-scale model of a jet engine fan. The average path loss was calculated, and the shapes of the fading distributions were extracted. The time between two consecutive fading dips was measured in the empirical part. Measurements were performed with engine speeds of both 30 and 60 rpm and were shown to be linearly scalable from 60 rpm to full speed of 10 000 rpm. The results showed an average path loss of about 55 dB. When scaling the measurements from 60 rpm to full-speed rotation of about 10 000 rpm, the fading was so severe that the time slot between consecutive fading dips was limited to 290 μs.