Color Response Modification of Encapsulated Liquid Crystals Used in Rotating Disk Heat Transfer Studies

The liquid crystal thermography can be successfully used in both transient and steady-state heat transfer experiments with excellent spatial resolution and good accuracy. Although most of the past liquid crystal based heat transfer studies are reported in the stationary frame, measurements from the rotating frame of turbomachinery systems exist The main objective of the present investigation is to determine the influence of rotation on the color calibration of encapsulated liquid crystals sprayed on the flat surface of a rotating aluminum disk. The investigation is performed for a rotational speed range from 0 rpm to 7500 rpm using three different liquid crystal coatings displaying red at 30, 35 and 45° C, under stationary conditions. An immediate observation from the present study is that the color response of liquid crystals is strongly modified by the centrifugal acceleration of the rotating environment. It is consistently and repeatedly observed that the hue versus temperature curve is continuously shifted toward lower temperatures by increasing rotational speed. The relative shift of the display temperature of the green can be as high as 7°C at 7500 rpm when compared to the temperature of the green observed under stationary conditions. The present study shows that relative shift of the liquid crystal color has a well-defined functional dependency to rotational speed. The shift is linearly proportional to the centrifugal acceleration. It is interesting to note that the individual shift curves of the green for all three liquid crystal coatings collapse into a single curve when they are normalized with respect to their own stationary green values. When the color attribute is selected as “intensity” instead of “hue”, very similar shifts of the temperature corresponding to the intensity maximum value appearing around green is observed. An interpretation of the observed color shift is made from a thermodynamics energy balance point of view.

Performance Improvement Through Indexing of Turbine Airfoils: Part 1 — Experimental Investigation

This paper describes the results of a study to determine the performance improvements achievable by circumferentially indexing successive rows of turbine stator airfoils. An experimental / analytical investigation has been completed which indicates significant stage efficiency increases can be attained through application of this airfoil clocking concept. A series of tests was conducted at the National Aeronautics and Space Administration’s (NASA) Marshall Space Flight Center (MSFC) to experimentally investigate stator wake clocking effects on the performance of the Space Shuttle Main Engine Alternate Fuel Turbopump Turbine Test Article. Extensive time-accurate Computational Fluid Dynamics (CFD) simulations have been completed for the test configurations. The CFD results provide insight into the performance improvement mechanism. Part one of this paper describes details of the test facility, rig geometry, instrumentation, and aerodynamic operating parameters. Results of turbine testing at the aerodynamic design point are presented for six circumferential positions of the first stage stator, along with a description of the initial CFD analyses performed for the test article. It should be noted that first vane positions 1 and 6 produced identical first to second vane indexing. Results obtained from off-design testing of the “best” and “worst” stator clocking positions, and testing over a range of Reynolds numbers are also presented. Part two of this paper describes the numerical simulations performed in support of the experimental test program described in part one. Time-accurate Navier-Stokes flow analyses have been completed for the five different turbine stator positions tested. Details of the computational procedure and results are presented. Analysis results include predictions of instantaneous and time-average mid-span airfoil and turbine performance, as well as gas conditions throughout the flow field. An initial understanding of the turbine performance improvement mechanism is described.

Feasibility Study on Oil Droplet Flow Investigations Inside Aero Engine Bearing Chambers: PDPA Techniques in Combination With Numerical Approaches

The present paper deals with oil droplet now phenomena in aero engine bearing chambers. An experimental investigation of droplet sizes and velocities utilizing a Phase Doppler Particle Analyzer (PDPA) has been performed for the first time in bearing chamber atmospheres under real engine conditions. Influences of high rotational speeds are discussed for individual droplet size classes. Although this is an important contribution to a better understanding of the droplet flow impact on secondary air/oil system performance, an analysis of the droplet flow behaviour requires an incorporation of numerical methods because detailed measurements as performed here suffer from both strong spatial limitations with respect to the optical accessibility in real engine applications and constraints due to the extremely time consuming nature of an experimental flow field analysis. Therefore, further analysis is based on numerical methods. Droplets characterized within the experiments are exposed to the flow field of the gaseous phase predicted by use of our well-known CFD code EPOS. The droplet trajectories and velocities are calculated within a Lagrangian frame of reference by forward numerical integration of the particle momentum equation. This paper has been initiated rather to show a successful method of bearing chamber droplet flow analysis by a combination of droplet sizing techniques and numerical approaches than to present field values as a function of all operating parameters. However, a first insight into the complex droplet flow phenomena is given and specific problems in bearing chamber heat transfer are related to the droplet flow.

Navier-Stokes and Potential Calculations of Axial Spacing Effect on Vortical and Potential Disturbances and Gust Response in an Axial Compressor

The effect of blade row axial spacing on vortical and potential disturbances and gust response is studied for a compressor stator/rotor configuration near design and at high loadings using 2D incompressible Navier-Stokes and potential codes, both written for multistage calculations. First, vortical and potential disturbances downstream of the isolated stator in the moving frame are defined; these disturbances exclude blade row interaction effects. Then, vortical and potential disturbances for the stator/rotor configuration are calculated for axial gaps of 10%, 20%, and 30% chord. Results show that the potential disturbance is uncoupled; the potential disturbance calculated from the isolated stator configuration is a good approximation for that from the stator/rotor configuration for all three axial gaps. The vortical disturbance depends strongly on blade row interactions. Low order modes of vortical disturbance are of substantial magnitude and decay much more slowly downstream than do those of potential disturbance. Vortical disturbance decays linearly with increasing mode except very close to the stator trailing edge. For a small axial gap, lower order modes of both vortical and potential disturbances must be included to determine the rotor gust response.

Numerical and Experimental Analysis of Secondary Flow in Modern State-of-the-Art Low Pressure Guide Vane Rows

To enhance the performance of the inlet guide vane and the annular duct of a jet engine, a detailed investigation of annular cascades with two different types of turbine guide vane rows is made. The first one is a leaned guide vane with an aspect ratio of two and a half and a transition duct ahead of the vane. To avoid the losses associated to the decelerating transition duct an alternative vane is designed and investigated with the same inlet and exit conditions. In this case the chord of the vane is increased to the effect that the vane begins immediately at the enterance of the diverging annulus and so a continuously accelerated flow is achieved. To maintain a good performance for this configuration a bowed-type vane with an aspect ratio of one is designed. The aim of the investigation is to obtain detailed informations on the secondary flow behaviour with particular regard to the development of the total pressure losses and the streamwise vorticity of the vortices inside and behind the blade rows. In the first step a three-dimensional, structured, explicit finite-volume flow-solver with a k–ε turbulence model is validated against the measurements, which were made in cross-sections behind the blades. Having proved that the numerical results are very close to the experimental ones, the secondary flow behaviour inside and behind the blade rows is analysed in the second step. By calculating the streamwise vorticity from the numerical results the formation of horse-shoe vortex, passage-vortex and the trailing edge vortex shed is investigated. The differences of the vortical motion and the formation of the total pressure losses between the two configurations of turbine guide vane rows are discussed.

Performance Assessment of an Annular S-Shaped Duct

An experimental investigation has been carried out to determine the aerodynamic performance of an annular S-shaped duct representative of that used to connect the compressor spools of aircraft gas turbine engines. For inlet conditions in which boundary layers are developed along an upstream entry length the static pressure, shear stress and velocity distributions are presented. The data shows that as a result of flow curvature significant streamwise pressure gradients exist within the duct, with this curvature also affecting the generation and suppression of turbulence. The stagnation pressure loss within the duct is also assessed and is consistent with the measured distributions of shear stress. More engine representative conditions are provided by locating a single stage compressor at inlet to the duct. Relative to the naturally developed inlet conditions the flow within the duct is less likely to separate, but mixing out of the compressor blade wakes increases the measured duct loss. With both types of inlet conditions the effect of a radial strut, such as that used for carrying loads and engine services, is also described both in terms of the static pressure distribution along the strut and its contribution to overall loss.

Theoretical Modelling of Relative Wall Motion Effects in Tip Leakage Flow

The influence of relative wall motion in modifying the leakage flow through the tip clearance is investigated. A theoretical model is developed, in order to calculate the mass flow rate through the gap. The physical mechanism by which relative wall motion affects the leakage flow is analyzed and the differences between the turbine and compressor case are identified. This model, being an extension of an already existing one, not taking into account relative wall motion, is incorporated into the tip clearance calculation procedure, already developed by the authors. Theoretical results of the complete calculation procedure (secondary flow plus tip clearance model) are compared with experimental data, for the case of compressor and turbine cascades, as well as for the case of a single rotor. Good agreement between theory and experiment is obtained.

Numerical Analysis of Fluid Flow in an Internal Turbine Blade Cooling Passage

The need to develop ultra-high efficiency turbines demands the exploration of methods which will improve the thermal efficiency and the specific thrust of the engine. One means of achieving these goals is to increase the turbine inlet temperature. In order to accomplish this, further advances in turbine blade cooling technology will have to be realized. A technique which has only recently been used in the analysis of turbine blade cooling is computational fluid dynamics. The purpose of this paper is to present a numerical study of the flowfield inside of the internal cooling passage of a radial turbine blade. The passage is modeled as two-dimensional and non-rotating. The flowfield solutions are obtained using a pseudo-compressible formulation of the Navier-Stokes equations. The spatial discretization is performed using a symmetric second-order accurate TVD (Total Variational Diminishing) scheme. Calculations are performed on a multi-block-structured grid. Turbulence is modeled using a modified κ-ω model. In the absence of experimental data, results appear to be realistic based on common experiences with fluid mechanics.

Numerical and Experimental Investigation of Secondary Flow in a High Speed Compressor Stator

A series of numerical and experimental studies have been conducted to understand the mechanism of loss generation in a high speed compressor stator with inlet radial shear flow over the span. In this study, numerical simulation is extensively used to investigate the complex three-dimensional flow in the cascades and to interpret the phenomena appeared in the high speed compressor tests. It has been shown that the inlet radial shear flow generated by upstream rotor had a significant influence on the stator secondary flow, and consequently on the total pressure loss. Redesign of the stator aiming at the reduction of loss by controlling secondary flow has been carried out and the resultant performance recovery was successfully demonstrated both numerically and experimentally.

Experimental Study of the Unsteady Boundary Layer Behavior on a Turbine Cascade

The unsteady boundary layer behavior on a turbine cascade is experimentally investigated and the results are presented in this paper. To perform a detailed study on unsteady cascade aerodynamics and heat transfer, a new large-scale, high-subsonic research facility for simulating the periodic unsteady flow has been developed. It is capable of sequentially generating up to four different unsteady inlet flow conditions that lead to four different passing frequencies, wake structures, and freestream turbulence intensities. For a given Reynolds number, three different unsteady wake formations are utilized. Detailed unsteady boundary layer velocity, turbulence intensity, and pressure measurements are performed along the suction and pressure surfaces of one blade. The results presented in the temporal-spatial domain display the transition and further development of the boundary layer, specifically the ensemble-averaged velocity and turbulence intensity.