Pre-Test Evaluation of Probe Configuration Errors

Rake probes are commonly used in turbomachinery applications to measure distorted inlet flows, including both pressure and swirl distortions. An especially common configuration is the 40 probe rake array, consisting of eight identical arms equally spaced circumferentially around the inlet, each arm having five sensing locations spaced along centers of equal area. By measuring many locations simultaneously, rake probes have the advantage of expedited data collection when compared to traversing probes. However, this reduction in test time comes at the cost of rigid geometry with limited measurement locations. As a result, it is possible for rake probes to miss or exaggerate significant areas of the flow profile, such as large gradients or small features, based on the fixed location of the probe and the particular details of the distortion. The purpose of this paper is to demonstrate a procedure that can be used to evaluate the ability of any desired probe configuration (40 probe rake or otherwise) to sufficiently and accurately measure a non-uniform flow profile. Results of this procedure for a range of profiles and probe configurations are also presented. In order to accurately determine the impact of discrete sampling on the results, two broad sets of data were generated numerically and analyzed. The first set consists of four fundamental total pressure distortions: once-per-rev circumferential, twice-per-rev circumferential, hub radial, and tip radial. The second set consists of three realistic turbofan distortion patterns: two analytic (though not fundamental) profiles, and one generated from S-duct computational results. For all investigated patterns, Radial Distortion Intensity and Circumferential Distortion Intensity are calculated in the manner described by ARP 1420, a guideline issued by the S-16 Turbine Engine Inlet Flow Distortion Committee for measuring total pressure distortions in turbomachinery. Additionally, interpolated total pressure contours are generated for each measurement configuration. These were then used to make point-to-point comparisons between the actual and estimated data. While total pressure distortion was used as the variable of interest for the majority of this paper, the conclusions may be applied to swirl, temperature, or any other flow property measured using a probe rake or traversing probe.

Physics-Based Simulation in Support of a Through-Life Gas Turbine Service Business Model

The overall aim of the work reported in this paper is to explore whether a physics-based simulation approach has the potential to reduce the uncertainty & variability associated with both predicting & managing maintenance costs and improving engine design to optimise through-life economic performance. The main novelty in the paper is to demonstrate how an innovative Digital Geometry model can represent typical in-service component degradation and then support appropriate simulation meshes to permit degraded performance to be predicted. Two examples are given: blade erosion from particulates; and a simulated cooled blade burn-through event.

Design and Maintenance Issues of Being Able to Repair Marine Gas Turbines Without Removal From the Ship

When the concept of aircraft derivative marine gas turbines were originally proposed, one of the selling points was the engine was going to be easy to remove and replace thereby minimizing the operational impact on the ship. Anticipated Mean Time Between Removal (MTBR) of these engines was expected to be approximately 3000 hours, due mostly to turbine corrosion damage. This drove the design and construction of elaborate removal routes into the engine intakes; the expected time to remove and replace the engine was expected to be less than five days. However, when the first USN gas turbine destroyers started operating, it was discovered that turbine corrosion damage was not the problem that drove engine maintenance. The issues that drove engine maintenance were the accessories, the compressor, combustors and engine vibration. Turbine corrosion was discovered to be a longer term affect. This was primarily due to the turbine blade and vane coatings used and intake air filtration. This paper discusses how engine design, tooling development, maintenance procedure development and engine design improvements all contributed to extending the MTBR of USN propulsion and electrical power generation gas turbines on the DD 963, CG 47, DDG 51 and FFG 7 classes to greater than 20,000 hours. The ability to remove the gas turbine rapidly or in most cases repair the engine in-place has given the USN great maintenance flexibility, been very cost effective and not impacted operational readiness.

Load Independent Losses of an Aeroengine Epicyclic Power Gear Train: Numerical Investigation

The development of Ultra-High Bypass Ratio (up to 20) engines with the aim of improving the propulsive efficiency, introduces new challenges for the transmission system in terms of heat management and power losses, since the amount of power transferred through the gearbox is greatly increased. In this respect the accurate estimate of losses at the various flow regimes realized during a typical aeroengine mission within a Power Gear Box (PGB) is essential for the correct design and operation of the engine itself. This paper proposes a computational methodology to estimate all fluid-dynamic (load-independent) losses, which become a major source of dissipation at the high rotational speeds typical of aeroengines, developing within an epicyclic PGB. The overall procedure is based on the superposition principle and approaches the three fluid-dynamic losses, namely injection, windage and meshing losses, with different numerical techniques. The simulations of windage effects, which consider the actual PGB geometry including the carrier disk, the lubricant spray-bar and the external casing, are based on steady-state computations. In order to introduce such simplification, a scaling procedure that avoids interference of the stationary and rotating interfaces was implemented following the outcomes of a previous analysis. The computation of meshing losses employs a fully unsteady dynamic mesh approach and considers a portion of the meshing gears only. Both the sun-planet and planet-ring meshing were considered showing that the latter introduces a much lower level of losses. Finally the injection losses are calculated considering the oil jet momentum variation with simplified methods based on 0D modelling. The proposed procedure, based on the superposition principle and applied to a planetary power gear train, is tested against experimental results described in a previous paper focused on a meshing gear pair.

Experimental Study of Flow Inside a Centrifugal Fan Using Magnetic Resonance Velocimetry

Three-dimensional, three-component time-averaged velocity fields have been measured within a low-speed centrifugal fan with forward curved blades. The model investigated is representative of fans commonly used in automotive HVAC applications. The flow was analyzed at two Reynolds numbers for the same ratio of blade rotational speed to outlet flow velocity. The flow patterns inside the volute were found to have weak sensitivity to Reynolds number. A pair of counter-rotating vortices evolve circumferentially within the volute with positive and negative helicity in the upper and lower regions, respectively. Measurements have been further extended to capture phase-resolved flow features by synchronizing the data acquisition with the blade passing frequency. The mean flow field through each blade passage is presented including the jet-wake structure extending from the blade and the separation zone on the suction side of the blade leading edge.

Updating the Helicopter Gas Turbine Engine Mass Model for the Defining of the Turboshaft Engine Optimal Operating Cycle Parameters

The article describes the adjusted parametrical turboshaft gas turbine engine mass model that is applied for the helicopter engine operating cycle parameters optimization during a conceptual engineering. During the operation of the take-off mass, which indirectly characterizes the cost of materials for the entire designed aircraft system, one of the main components which determines the coordination of the helicopter and its engine parameters is a mass of the gas turbine power unit. Moreover, during the parametrical studies the designed mass of a power unit should be defined by the parameters of a gas turbine engine; however, this type of dependencies is not that well enough studied for today. Therefore the evaluation of the dependency between the engine mass and its operational parameters is performed by using either generalized statistical data for existing designs or by parametrical mass models since there is nothing more precise up to date. However as new types of gas turbine engines appear it is required to update the values of parametrical model coefficients. This article describes the influence of different cooling system units on the engine mass and also clarifies the coefficients that specify the engine mass advance by introducing the structural-technological measures. The last one is highly dependent on the designed gas turbine engine (GTE) serial production year. It also has been proposed to represent some coefficients that are used in the model as dependencies of the main operational parameters. This has allowed to perform the parametrical study and to gain predictive solutions in correspondence to the modern engine design level.

CFD and Experimental Analysis of Centrifugal Fans With Forward Curved Blades Used in Electric Motors

This work presents an analysis of the aerodynamic performance of a centrifugal fan with forward curved blades (Sirocco) applied to electric motors. In this analysis were carried out computational fluid dynamics (CFD) simulations and experimental tests for comparison of results. The focus of this analysis is the performance comparison among three different models of general connection interface that are required for the connection between the grids of the rotating and stationary domains of CFD simulation, considering the method adopted by the Ansys CFX, software used as computational tool. Thereby, Frozen Rotor, Stage, and Transient Rotor-Stator were the interface models evaluated. For comparison reference, the experimental data were used to evaluate the performance of each interface models for overall operating range of the fan.

Strongly Coupled Fluid-Structure Interaction Simulation of a 3D Printed Fan Rotor

Additive manufacturing represents a new frontier in the design and production of rotor machines. This technology drives the engineering research framework to new possibilities of design and testing of new prototypes, reducing costs and time. On the other hand, the fast additive manufacturing implies the use of plastic and light materials (as PLA or ABS), often including a certain level of anisotropy due to the layered deposition. These two aspects are critical, because the aero-elastic coupling and flow induced vibrations are not negligible for high aspect ratio rotors. In this work, we investigate the aeroelastic response of a small sample fan blade, printed using PLA material. Scope of the work is to study both the structure and flow field dynamics, where strong coupling is considered on the simulation. We test the blade in two operating points, to see the aero-mechanical dynamics of the system in stall and normal operating condition. The computational fluid-structure interaction (FSI) technique is applied to simulate the coupled dynamics. The FSI solver is developed on the base of the finite element stabilized formulations proposed by Tezduyar et al. We use the ALE formulation of RBVMS-SUPS equations for the aerodynamics, the non-linear elasticity is solved with the Updated Lagrangian formulation of the equations of motion for the elastic solid. The strong coupling is made with a block-iterative algorithm, including the Jacobian based stiffness method for the mesh motion.

A New Practical Approach to the Design of Industrial Axial Fans: Part II — Forward-Swept Blades With Low Hub-to-Tip Ratio

The authors previously suggested a simple method to design forward-swept axial-flow rotors with blades having low hub-to-tip and high aspect ratios. This design method was demonstrated experimentally to increase the aeraulic performance of a small tube-axial fan having unswept blades and 0.4 hub-to-tip ratio, while maintaining the efficiency in the entire operation range. However, the method has not yet been assessed by experimental tests of lower hub-to-tip ratio designs where the strong three-dimensionality of the actual blade passage flow could compromise its validity. This assessment is the object of the present paper, which is aimed at examining the practical effectiveness of the forward-swept blade design method for low hub-to-tip ratio tube-axial fans. To this end, past results of the authors’ work are supported here by the design of a new 315mm forward-swept industrial fan derived from the 0.28 hub-to-tip ratio design presented in Part I of this paper. The ISO-5801 aerodynamic performance tests at blade Reynolds number of approximately 60,000 show that the method permits the design of forward-swept industrial fans capable of pressure coefficients in excess of 0.02 at aeraulic efficiency well above 60%, in a wide range of flow rate coefficients and blade positioning angles. Moreover, the method allows obtaining a pressure coefficient equal to 0.021 at 70% maximum efficiency, with an improvement of both the stall margin and stable operation pressure curve of the unswept design, if applied in combination with the complete fan design method presented in Part I of this paper.

Using Machine Learning to Predict Core Sizes of High-Efficiency Turbofan Engines

With the rise in big data and analytics, machine learning is transforming many industries. It is being increasingly employed to solve a wide range of complex problems, producing autonomous systems that support human decision-making. For the aircraft engine industry, machine learning of historical and existing engine data could provide insights that help drive for better engine design. This work explored the application of machine learning to engine preliminary design. Engine core-size prediction was chosen for the first study because of its relative simplicity in terms of number of input variables required (only three). Specifically, machine-learning predictive tools were developed for turbofan engine core-size prediction, using publicly available data of two hundred manufactured engines and engines that were studied previously in NASA aeronautics projects. The prediction results of these models show that, by bringing together big data, robust machine-learning algorithms and automation, a machine learning-based predictive model can be an effective tool for turbofan engine core-size prediction. The promising results of this first study paves the way for further exploration of the use of machine learning for aircraft engine preliminary design.