Exit Plane Research Articles

Investigation on Jet Characteristics of Turbofan Exhaust System under Take-off Condition with High Angle of Attack

To understand the jet flow characteristics of turbofan separate exhaust system, a parametric design method based on the initial Class Shape Transformation function was developed. HBPR and UHBPR turbofan separate exhaust systems were designed. Furthermore, the jet flow characteristics of the HBPR turbofan exhaust system under take-off condition with zero angle of attack were studied based on numerical simulation. The jet flow characteristics of the HBPR and UHBPR turbofan exhaust system under take-off condition with high angle of attack were also simulated. The effects of angle of attack and bypass ratio on the jet flow characteristics were investigated and the related flow mechanisms were analyzed. Results show that the axisymmetric plumes of the HBPR turbofan exhaust system are distributed around the engine axis under take-off condition with zero angle of attack. With the plug wake as the center, the core flow, the fan/core shear layer, the fan flow, the fan/free stream shear layer and the free stream are wrapped around the plug wake from inside out. Vortexes appear in the lee area at the back of the cowl and jet flow under take-off condition with high angle of attack. These vortexes cause cross sectional secondary flow and expose the high-velocity core flow to the low-velocity free stream. The contact area and velocity gradient in the mixing region among the free stream, fan flow and core flow increase. Therefore, the mixture among jet flow and free stream strengthens. So the high-velocity region, the high-vorticity region, and the high turbulence kinetic energy region shorten by 55.1%, 47.7% and 50.9% respectively. The vorticity values and turbulence kinetic energy level peak on the upper side of the exhaust plumes increase by about 30% and 87% respectively. Relative to these parameters from the HBPR turbofan exhaust system, the jet velocity peak value of UHBPR turbofan decreases by 5.5% under take-off condition with high angle of attack. The vorticity values and turbulence kinetic energy level reduce due to decreased velocity gradient in shear layers downstream of the nozzle exit plane. The turbulence kinetic energy level peak on the upper side of the exhaust plumes decreases by 29.3%. The reasons are that the contact area between high-velocity core flow and the free stream decreases due to thicker fan flow and the velocity gradient in the core flow and free stream mixing region decreases because of the lower core flow velocity.

Inlet Flow Distortion in an Advanced Civil Transport Boundary Layer Ingesting Engine Installation

AbstractThis article presents first-of-a-kind measurements, and complementary computations, of the flow through the propulsion system of a boundary layer ingesting, twin-engine advanced civil transport aircraft configuration. The experiments were carried out in the NASA Langley 14- by 22-foot Subsonic Tunnel, using a 1:11 scale model of the D8 “double-bubble” aircraft with electric ducted fans providing propulsive power. Overall force and moment measurements and flow field surveys at the inlet and nozzle exit planes were obtained. The computations were carried out with the NASA OVERFLOW code. The measurements and computations were conducted for a range of aircraft angles of attack and propulsor powers representing operating points during the aircraft mission. Velocity and pressure distributions at the propulsor inlet and exit, and integral inlet distortion metrics, are presented to quantify the flow nonuniformity due to boundary layer ingestion. The distorted inflow exhibits qualitative and quantitative changes over the mission, from a unidirectional stratified stagnation pressure at cruise to a streamwise vortex structure at climb conditions. The computations capture these flow features and reveal the interactions between airframe and propulsor that create these three-dimensional flow variations.

Numerical Optimization of a Premixer for an Internal Combustion Engine using Producer Gas as a Fuel

Gasification seems to be one of the sustainable green energy solutions to fulfill the current and future energy needs. For efficient utilization of producer gas on existing IC Engines, carburetor/premixer needs to be carefully designed and developed to achieve uniform mixing quality. A long radius nozzle type premixer has been designed for natural gas engine to be operated on producer gas as an alternate fuel. Different configurations of T – Type premixers with single air entry and twin air entry with different throat diameters and hole sizes are numerically analysed using ANSYS® CFX. Turbulence is modelled using RNG k - ε closure model. Mixer performance is compared in terms of constituents’ mass fraction, flow Uniformity Index (UI) and pressure penalty. Numerical analysis reveals that throat diameter, air entry type and air hole diameter governs mixing and pressure drop. Out of all configurations, twin air entry type premixer provides better mixing of producer gas and air. The optimized design of premixer shows that the absolute deviation in mass fraction of individual constituent lies in the range of ± 1.73% with respect to the actual mass fractions obtained. The average absolute deviation calculated is 1.37% with Uniformity Index 0.958 at the exit plane while the pressure drop across the premixer is 951 Pa.

Effect of profiled endwall on heat transfer under different turbulence intensitie

The efficiency of the gas turbine engine has been enhanced by improving aerodynamic performance and reducing heat load through control of secondary vortices according to the first-stage vane endwall configurations. In this study, the heat transfer features of flat and profiled endwalls were investigated under different turbulence intensities. Experiments were conducted under a fixed turbine-vane exit Reynolds number of 300,000 and Mach number of 0.15 on both geometries. The test specimen was scaled up 3.23 times based on the actual gas turbine vane geometry. The profiled endwall configuration was optimized considering the combustor outlet/exit geometry. Five control points were set for the vane geometry, and the minimum aerodynamic loss at the vane exit plane was determined as an objective function. As the turbulence intensity is increased, the heat transfer is increased on both endwall geometries. However, the intensity of the horseshoe vortex was significantly reduced with the profiled endwall compared with the flat endwall. Consequently, the area-averaged Sherwood number on the profiled endwall was about 25.8% (under the low turbulence intensity) and 19% (under the high turbulence intensity) lower than those of a flat endwall case, respectively.

Evidence of a free-space ion acceleration layer in the plume of a quad confinement plasma source

The quad confinement plasma source is a novel plasma device developed for space propulsion applications, whose core is an E×B discharge with open electron drift. The magnetic field is produced by independently powered electromagnets able to generate different magnetic field topologies with the ultimate aim of manipulating the ion flow field for achieving thrust vectoring. In this work, we map the ion velocity in the plasma ejected from the quad confinement thruster with different magnetic configurations using non-intrusive laser-induced fluorescence diagnostics. Measurements show a steep ion acceleration layer located 8 cm downstream the exit plane of the discharge channel, detached from any physical boundary of the plasma source. In this location, the ion velocity increases from 3 to 10 km/s within a 1 cm axial region. The ion acceleration profile has been characterized under multiple testing conditions in order to identify the influence of the magnetic field intensity and topology on this peculiar ion acceleration layer.

Three-dimensional tomographically reconstructed optical emission profiles of Hall thruster plasmas

A diagnostic system was developed for spectrally resolved, three-dimensional tomographic reconstruction of Hall thruster plasmas, and local intensity profiles of Xe I and Xe II emissions were reconstructed. In this diagnostic system, 28 virtual cameras were generated using a single, fixed charge-coupled device camera by rotating the Hall thruster to form a sufficient number of lines of sight. The Phillips–Tikhonov regularization algorithm was used to reconstruct local emission profiles from the line-integrated emission signals. The reconstruction performance was evaluated using both azimuthally symmetric and asymmetric synthetic phantom images including 5% Gaussian white noise, which resulted in a root-mean-square error of the reconstruction within an order of 10−3 even for a 1% difference in the azimuthal intensity distribution. Using the developed system, three-dimensional local profiles of Xe II emission (541.9 nm) from radiative decay of the excited state 5p4(3P2)6p2[3]°5/2 and Xe I emission (881.9 nm) from 5p5(2P°3/2)6p2[5/2]3 were obtained, and two different shapes were found depending on the wavelength and the distance from the thruster exit plane. In particular, a stretched central jet structure was distinctively observed in the Xe II emission profile beyond 10 mm from the thruster exit, while gradual broadening was found in the Xe I emission. Approximately 10% azimuthal nonuniformities were observed in the local Xe I and Xe II intensity profiles in the near-plume region (<10 mm), which could not be quantitatively distinguished by analysis of the frontal photographic image. Three-dimensional Xe I and Xe II intensity profiles were also obtained in the plume region, and the differences in the structures of both emissions were visually confirmed.

CFD Analysis of S-Duct with Varying Area Distributions

The ideal design of an intake S-duct (serpentine duct) of the jet engine should be able to supply adequate air regardless of the operating conditions of the aircraft, efficiently decrease the speed of incoming compressible flow and turn it to the desired location. Total pressure losses should be as low as possible and pressure recovery should be as high as possible. To maximize the efficiency, the vortex generated due to the two curvatures must be suppressed. This can be achieved by varying the area distribution of the duct which helps in isolating the effects of diffusion and curvature as much as possible. In this numerical analysis, six different serpentine intake ducts are designed with varying area distributions, but a constant centerline. The inlet area and the exit area are kept constant along with the centerline. Total pressure contours and transverse velocity contours are observed for flow in each duct to find the duct configuration with least losses. The geometry of a duct with least total pressure loss is considered as the duct with optimum area distribution. The inlet boundary condition is set as the velocity inlet condition with velocity 204.1m/s while, the outlet boundary condition is set as the pressure outlet condition, with outlet pressure of 1 atm. The losses are found to be minimal in the ‘duct 5’ with least transverse flow. This is due to the efficient area distribution which result in initial acceleration and completion of diffusion much before the exit plane causing suppression of vortex at the bend. This in turn helps the ‘duct 5’ to retain its flow energy along the centerline direction resulting in better pressure recovery.

A Nonhomogeneous Bulk Flow Model for Gas in Liquid Flow Annular Seals: An Effort to Produce Engineering Results

Abstract Seals in multiple-phase rotordynamic pumps must operate without compromising system efficiency and stability. Both field operation and laboratory experiments show that seals supplied with a gas in liquid mixture (bubbly flow) can produce rotordynamic instability and excessive rotor vibrations. This paper advances a nonhomogeneous bulk flow model (NHBFM) for the prediction of the leakage and dynamic force coefficients of uniform clearance annular seals lubricated with gas in liquid mixtures. Compared to a homogeneous bulk flow model (HBFM), the current model includes diffusion coefficients in the momentum transport equations and a field equation for the transport of the gas volume fraction (GVF). Published experimental leakage and dynamic force coefficients for two seals supplied with an air in oil mixture whose GVF varies from 0% (pure liquid) to 20% serve to validate the novel model as well as to benchmark it against predictions from a HBFM. The first seal withstands a large pressure drop (∼38 bar) and the shaft speed equals 7.5 krpm. The second seal restricts a small pressure drop (1.6 bar) as the shaft turns at 3.5 krpm. The first seal is typical as a balance piston whereas the second seal is found as a neck-ring seal in an impeller. For the high pressure seal and inlet GVF = 0.1, the flow is mostly homogeneous as the maximum diffusion velocity at the seal exit plane is just ∼0.1% of the liquid flow velocity. Thus, both the NHBFM and HBFM predict similar flow fields, leakage (mass flow rate), and drag torque. The difference between the predicted leakage and measurement is less than 5%. The NHBFM direct stiffness (K) agrees with the experimental results and reduces faster with inlet GVF than the HBFM K. Both direct damping (C) and cross-coupled stiffness (k) increase with inlet GVF &amp;lt; 0.1. Compared to the test data, the two models generally under predict C and k by ∼25%. Both models deliver a whirl frequency ratio (fw) ∼ 0.3 for the pure liquid seal, hence closely matching the test data. fw raises to ∼0.35 as the GVF approaches 0.1. For the low-pressure seal, the flow is laminar; the experimental results and both NHBFM and HBFM predict a null direct stiffness (K). At an inlet GVF = 0.2, the NHBFM predicted added mass (M) is ∼30% below the experimental result while the HBFM predicts a null M. C and k predicted by both models are within the uncertainty of the experimental results. For operation with either a pure liquid or a mixture (GVF = 0.2), both models deliver fw = 0.5 and equal to the experimental finding. The comparisons of predictions against experimental data demonstrate that the NHBFM offers a marked improvement, in particular for the direct stiffness (K). The predictions reveal that the fluid flow maintains the homogeneous character known at the inlet condition.

Heat transfer due to turbulent annular impinging jet with a bullet extension at the end of the inner blockage rod

The flow and thermal behavior of a turbulent annular jet impinging on a flat plane are numerically investigated to examine the effect of extending the inner rod of the annulus beyond the exit plane of the jet. The Fluent commercial code is used for conducting the computations employing the k-ε (Realizable) turbulence model. Three different shapes of the extended annular rod end, viz., cylindrical, conical, and semi-ellipsoidal, are considered. The length of the extended part is taken as 0.5, 1, and 2 times the outer jet diameter. The extension of the annular rod end modifies the flow structure and eliminates the recirculation region at the exit of the annular jet, which noticeably affects the impingement plane heat transfer. Depending on the shape and length of the bullet, the bullet extensions affect the heat transfer differently. Compared to the case without the extension, the conical extension diminishes the heat transfer, the ellipsoidal extension moderately increases the heat transfer, while the cylindrical extension enhances the heat transfer for a larger bullet length (≥1). Also, the location and the magnitude of the peak Nusselt number at the hot impingement surface vary as the shape and length of the extension change.

Temperature distributions of internal flow and external jet fields of double serpentine convergent nozzle for turbofan

In order to clarify the temperature distribution of a double-serpentine convergent nozzle wall and the exhaust plume characteristics downstream the nozzle exit based on the real exhaust mixer configuration of a turbofan engine, the temperature distribution of the internal and external jet flow fields of the double-serpentine convergent nozzle are simulated numerically. The simulation results show that the coupling effect between S-shaped circular-to-rectangular profile and the exhaust mixer has a joint influence on the "hot streak" distribution of the double-serpentine convergent nozzle wall and the exhaust plume characteristics downstream the nozzle exit. The "Hot streak" is distributed in the shape similar to several ribbons in the central areas of the upper and lower wall of the second serpentine passage and in the exit area. The maximum value of temperature on the lower wall is 867 K and increases 25.1% compared with that on the upper wall. However, the average value of temperature of the lower wall is merely 7.1% higher than that of the upper wall. The "whale tail" high-temperature core area is formed on the exit plane of the double-serpentine convergent nozzle. It leads to two wide high-temperature jet flow areas in the high-temperature core area downstream the nozzle exit. The low-temperature flow occurs along the two sides of the exhaust plume core area due to the lateral-direction motion towards vortices.

Time-synchronized laser-induced fluorescence in the near-field of a 600 Watt Hall thruster

We report on the results of an experimental campaign to measure time-varying velocity distributions in the near-field of a low power Hall thruster. We employ a sample-hold technique, enhanced by parallelizing the measurement hardware into several signal processing channels that vastly increases the data acquisition rate. The measurements are applied to study flow field dynamics in a commercial BHT-600 Hall thruster undergoing unforced breathing mode oscillations in the 44–49 kHz range. A very detailed experimental picture of the near-field emerges from these studies. The results indicate that velocity fluctuations lessen further downstream of the exit plane. Along the thruster axis where there is a general appearance of a central jet, there is evidence of a low velocity ion population in between the periodic bursts of high velocity ions, indicative of local ionization of neutrals outside of the thruster. One possible source of this residual ionization may be background chamber gas, which is not unexpected with the limited pumping capacity of ground test facilities.

Impact of a Centerbody on the Unsteady Flow Dynamics of a Swirl Nozzle: Intermittency of Precessing Vortex Core Oscillations

Abstract The precessing vortex core (PVC) is a self-excited flow oscillation state occurring in swirl nozzles. This is caused by the presence of a marginally unstable hydrodynamic helical mode that induces precession of the vortex breakdown bubble (VBB) around the flow axis. The PVC can impact emissions and thermoacoustic stability characteristics of combustors in various ways, as several prior studies have shown. In this paper, we examine the impact of centerbody diameter (Dc) on the PVC in a nonreacting flow in a single nozzle swirl combustor. Time-resolved high-speed stereoscopic PIV measurements are performed for combinations of two swirl numbers, S = 0.67 and 1.17 and Dc = 9.5 mm, 4.73 mm, and 0 (i.e., no centerbody). The bulk flow velocity at the nozzle exit plane is kept constant as Ub = 8 m/s for all cases (Re∼20,000). The centerbody end face lies in the nozzle exit plane. A new modal decomposition technique based on wavelet filtering and proper orthogonal decomposition provides insight into flow dynamics in terms of global modes extracted from the data. The results show that without a centerbody, a coherent PVC is present in the flow as expected. The introduction of a centerbody makes the PVC oscillations intermittent. These results suggest two routes to intermittency as follows. For S = 0.67, the VBB and centerbody wake recirculation zone regions are nominally distinct. Intermittent separation and merger due to turbulence result in PVC oscillations due to the destabilization of the hydrodynamic VBB precession mode of the flow. In the S = 1.17 case, the time averaged VBB position causes it to engulf the centerbody. In this case, the emergence of intermittent PVC oscillations is a result of the response of the flow to broadband stochastic forcing imposed on the time averaged vorticity field due to turbulence.

Design of Magnetic Circuit for Stationary Plasma Thruster

SPT-100 electrostatic thruster is considered, and the effects of magnetic circuit is studied by introducing magnetic screen. The magnetic flux density in the discharge channel is generated with the help of one inner coil and four outer coils. The radial magnetic field has to be maximum near the exit plane of the thruster to trap the electrons in acceleration region which are emitted from an external hollow cathode. These electrons help in increasing the ionization rate of the propellant gas. This is obtained by placing magnetic poles near exit plane. It helps to traps the electrons emitted from the external hollow cathode. The magnetic circuit should be designed such that the magnetic flux density is near to zero at the anode plane to reduce interaction of electrons with channel walls. To arrive at such better design, magnetic screens are used. Computational simulations are performed to quantify the magnetic flux density distribution along the channel using COMSOL Multiphysics software. The simulation results show that the obtained radial magnetic flux density is maximum near the exit plane, and the magnetic screens help in reducing the magnetic field at the anode region while maintaining the maximum magnetic field at the exit plane.

Vortex Formation Times in the Glottal Jet, Measured in a Scaled-Up Model.

In this paper, the timing of vortex formation on the glottal jet is studied using previously published velocity measurements of flow through a scaled-up model of the human vocal folds. The relative timing of the pulsatile glottal jet and the instability vortices are acoustically important since they determine the harmonic and broadband content of the voice signal. Glottis exit jet velocity time series were extracted from time-resolved planar DPIV measurements. These measurements were acquired at four glottal flow speeds (uSS = 16.1–38 cm/s) and four glottis open times (To = 5.67–23.7 s), providing a Reynolds number range Re = 4100–9700 and reduced vibration frequency f* = 0.01–0.06. Exit velocity waveforms showed temporal behavior on two time scales, one that correlates to the period of vibration and another characterized by short, sharp velocity peaks (which correlate to the passage of instability vortices through the glottis exit plane). The vortex formation time, estimated by computing the time difference between subsequent peaks, was shown to be not well-correlated from one vibration cycle to the next. The principal finding is that vortex formation time depends not only on cycle phase, but varies strongly with reduced frequency of vibration. In all cases, a strong high-frequency burst of vortex motion occurs near the end of the cycle, consistent with perceptual studies using synthesized speech.

100 keV vacuum sealed field emission gun for high resolution electron microscopy

A standalone 100 kV field emission gun (FEG) has been developed that can be installed and operated on a standard transmission electron microscopy electron optical column or custom designed high voltage electron optical columns. The FEG comprises a thermally assisted field emission cathode and an asymmetric electrostatic lens that can operate from 20 to 100 kV in an ultrahigh vacuum (UHV) chamber. In its current configuration, the FEG has spherical and chromatic aberration coefficients (Cs and Cc, respectively) in the range of Cs = 607–670 mm and Cc = 60–87 mm at 100 keV over a range of working distances of 50–206 mm from the exit plane of the FEG unit. A dedicated high voltage supply unit with voltage ripples of less than 1 ppm at 100 kV has also been developed. The FEG is transported under UHV and does not require the use of SF6 gas during operation, as is customary in high voltage FEG TEMs. Preliminary results of operating the FEG on a Philips Tecnai 12 and a JEOL JEM-1400HR TEM show the resolution of gold (111) crystal planes at 0.235 nm and (200) planes at 0.202 nm.

Four-Wall Flow Separation Control Using Microvortex Generators in Supersonic Duct

Suppression of shock-induced separation on all the walls of a rectangular duct using arrays of microvortex generators was experimentally investigated. The flowfield was analyzed using normal and inclined schlieren visualizations, surface oil flow visualization, wall static pressure distribution, and exit plane stagnation pressure distribution. The top wall and side wall array configurations were fixed, and only the bottom wall configuration was changed. It was found that control decreased the axial oscillations but increased the spanwise oscillations of the shock (oscillations in the spanwise length of the normal part of the shock). Among the control cases, as the control on the bottom wall increased, the exit plane stagnation pressure recovery, static pressure recovery across the shock, and area of the normal portion of the normal shock increased and then decreased. However, the attached flow percentage on the bottom wall increased with an increase in control. As the shock moved away from the microvortex generator arrays, the case with no control on any walls gave better results in terms of shock size and exit plane stagnation pressure recovery. Also, as the number of control devices on the bottom wall increased, the extent of flow separation on the bottom wall decreased; however, it had detrimental effects on the duct flow beyond some extent.

Rotating co-flow jets with horizontal lip and shape transition

Purpose Control over large-scale coherent structures and stream-wise vortices lead to enhanced entrainment/conservation of the jet which is desirable for most free jet applications such as design of combustion chamber in jet engines and flame length elongation of welding torch used for metal cutting. Design/methodology/approach A co-flow nozzle with lip thickness of 2 mm, between the primary (inner) and secondary (outer) flow, is selected. Three nozzle combinations are used, i.e. C–C (circle–circle), C–E (circle–ellipse) and C–S (circle–square) for acquiring comparative data. For these nozzle combinations, inner nozzle exit plane is kept as a circle, whereas the outer nozzle exit planes are varied to circle, ellipse and square. The exit plane area of outer nozzle for the nozzle combinations has equivalent diameter, De. The nozzles are fabricated in a way that the outer nozzle can be rotated along the longitudinal axis, keeping the inner nozzle intact. Findings The C–C nozzle combination is effective in low Mach number regime in decaying the jet, when the rotational component is introduced. Around 30% reduction in the jet core length is observed for the C–C nozzle combinations without any lip. The C–E nozzle shows sedative result in decaying or preserving the jet. The C–S nozzle combination shows interesting phenomenon, whereby the low subsonic case tends to conserve the jet by 15% and the higher subsonic case tends to decay the jet by 10%. Originality/value The developed nozzle systems show both conservative and destructive effect on the jet, which is desirable for the mentioned applications.

Spray characteristics of standard and alternative aviation fuels at high ambient pressure conditions

The spray characteristics of standard and alternative aviation fuels generated by a hybrid pressure-swirl airblast (HPSA) atomizer were investigated at high ambient pressure conditions. The measurements were performed in an optically accessible pressure vessel at pressures of 1, 2, 3, 4, 5, and 9.5 bar with heated gas and fuel temperatures of 394 K and 332 K, respectively. The drop size and drop velocity of sprays from a standard (Jet-A) and an alternative (C-5) aviation fuels were measured using phase Doppler anemometry (PDA), and 2-D visualization of spray was imaged using simultaneous fuel-Planar Laser Induced Fluorescence (PLIF) and Mie scattering. The spatially resolved PDA drop size measurements were obtained at 25.4 mm downstream of the swirler exit plane and up to 30 mm in the radial direction from the spray centerline. The Sauter Mean Diameter (D32) was observed to decrease significantly with increasing ambient pressure, up to 5 bar. However, the change in D32 was observed to be limited with a further increase in the ambient pressure. A higher swirler pressure drop resulted in a significant reduction in the D32 at an ambient pressure of 5 bar. Using simultaneous fuel-PLIF and Mie scattering imaging, the feasibility of the laser sheet drop-sizing (LSD) technique was examined for the spray at high ambient pressure application. In addition, the effect of ambient pressure on the spray cone angle was investigated using both instantaneous and averaged Mie images. The distribution of fuel vapor and droplets in the spray was also imaged and identified by comparing instantaneous fuel-PLIF and Mie images. A phenomenological three-step atomization model was used to predict the drop size and demonstrate the drop size trend with increasing ambient pressure.

Slit homogenizer introduced performance gain analysis based on the Sentinel-5/UVNS spectrometer

Abstract. Spatially heterogeneous Earth radiance scenes affect the atmospheric composition measurements of high-resolution Earth observation spectrometer missions. The scene heterogeneity creates a pseudo-random deformation of the instrument spectral response function (ISRF). The ISRF is the direct link between the forward radiative transfer model, used to retrieve the atmospheric state, and the spectra measured by the instrument. Hence, distortions of the ISRF owing to radiometric inhomogeneity of the imaged Earth scene will degrade the precision of the Level-2 retrievals. Therefore, the spectral requirements of an instrument are often parameterized in the knowledge of the ISRF over non-uniform scenes in terms of shape, centroid position of the spectral channel and the full width at half maximum (FWHM). The Sentinel-5/UVNS instrument is the first push-broom spectrometer that makes use of a concept referred to as a slit homogenizer (SH) for the mitigation of spatially non-uniform scenes. This is done by employing a spectrometer slit formed by two parallel mirrors scrambling the scene in the along track direction (ALT) and hence averaging the scene contrast only in the spectral direction. The flat mirrors do not affect imaging in the across track direction (ACT) and thus preserve the spatial information in that direction. The multiple reflections inside the SH act as coherent virtual light sources and the resulting interference pattern at the SH exit plane can be described by simulations using scalar diffraction theory. By homogenizing the slit illumination, the SH strongly modifies the spectrograph pupil illumination as a function of the input scene. In this work we investigate the impact and strength of the variations of the spectrograph pupil illumination for different scene cases and quantify the impact on the ISRF stability for different types of aberration present in the spectrograph optics.

Influence of inlet swirl on pattern factor and pressure loss in an aero engine combustor

The effect of compressor exit swirl angle (θsw) at the intake of an aero engine combustor on the exit temperature non-uniformity (pattern factor) and combustor total pressure loss is investigated. Experiments are conducted in the engine test rig, measuring the gas temperature and pressure at the inlet and exit planes of the combustor. These parameters are measured at distinct locations along the circumferential and radial directions in the engine test facility. Simulations are carried out using RANS based turbulence modeling and reacting flow approach with Ansys CFX commercial code. The predicted results are validated with experimental data at 5° swirl angle. The swirl angle at the combustor intake is further varied from 0° to 15° and 4 cases (0°, 5°, 10° and 15°) has been considered to predict the effect on the combustor pattern factor and pressure loss. The changes in the flow structure inside the combustion chamber for all these 4 cases are reported in detail. The pattern factor varies from 0.34 to 0.49 as swirl angle changes from 0° to 15°. The lowest pattern factor of 0.34 occurs at 10° swirl angle. However a linear increase in combustor total pressure loss from 5.85% to 6.53% is predicted with the change in swirl angle from 0° to 15°.