Proceedings of the National Aerospace Propulsion Conference

The present study intends to improve the performance of transonic axial flow compressor stage and its operating range by implementing passive flow control technique, a Casing Treatment. A plenum chamber with two different volumes placed above the bend skewed slots was implemented. The objective was to understand the effect of plenum volume on the performance of transonic axial flow compressor retrofitted with bend skewed casing treatment. The porosity of the selected bend skewed casing treatment was 33%. A detailed steady-state CFD analysis has been carried out for the compressor operating at six different speeds. Axial location of the casing treatment above the rotor tip was chosen based upon the previous experiments reported in literature [1]. For the same axial location and porosity, plenum chamber depth was varied from zero depth to full plenum depth to understand effect of a plenum volume. The results were compared with baseline model with solid casing wall. Significant improvement in stall margin was observed at all rotational speeds. Minor deviation was observed in stage total pressure ratio with reduction in efficiency at design speed.

The sensitivity of the fitness function comprising of weight coefficients assigned to performance variables in a genetic algorithm for meanline design of a transonic compressor is studied. The sum of the weight coefficients is unity. Six performance variables considered are the pressure ratio, efficiency, De-Haller numbers (for rotor and stator), and diffusion factors (for rotor and stator). Based on prior trials, the optimum weight coefficients for pressure ratio and efficiency were considered 0.3 each in the fitness function. Hence the sum of the weight coefficients for the two De-Haller Numbers and two Diffusion Factors considered is 0.4. The values of assigned weights have a significant impact on optimization outcome. Optimized design trials of weight coefficients with higher weightage to DFR resulted in higher efficiency with lower pressure ratio. Optimized design trials with higher weightages to DEHR and DEHS yielded into higher pressure ratio but lower efficiency. The data generated provides a guideline to choose combinations of weight coefficients for fitness functions for several performance requirements of a similar class of compressors for various applications.

Aeroelastic flutter investigation of a three-stage transonic axial compressor rotor is discussed in this paper. Unsteady CFD analyses were used to evaluate the flutter instability of the test compressor. Investigation on the effect of inlet guide vane and blade stiffness on aeroelastic damping is the prime objective of this study. The blade was subjected to vibration in CFD domain at specified frequency and amplitude. The unsteady aerodynamic force and work done by the blade at each vibration cycle were evaluated using fluid structure interaction technique. Energy method and work per cycle approach were adapted for this flutter prediction. Based on the work per cycle value, aerodynamic damping ratio was evaluated. A computational framework has been developed to calculate work per cycle and thereby aerodynamic damping ratio. Based on the magnitude and sign of aerodynamic damping ratio, occurrence of flutter was evaluated at all operating conditions of the engine. The primary cause for blade flutter was identified as large flow separation and flow unsteadiness due to high incidence on blade suction surface. The flow unsteadiness resulted into aerodynamic load fluctuation which matched with blade natural frequency and further resulted in blade excitation. Flutter boundary was evaluated for both with and without inlet guide vane case. Significant improvement in flow pattern and flutter boundary was observed for the case with inlet guide vanes.

Failure analysis and defect investigation of aero gas turbine engines, which are complex and safety-critical systems, call for advanced tools and techniques in the customer perspective framework of Strategic Quality Management (SQM). Functional Resonance Analysis Method (FRAM) is such a tool for failure mode and root cause analyses, helpful in instituting suitable corrective and preventive actions. Repeated failure of the drive shaft of oil cooling system of a turboshaft aero-engine at the Shear neck resulted in many premature withdrawals affecting fleet serviceability. Defect investigations attributed the failure initially to the Nitrided surface of the drive shaft, then the Heat treatment process, and subsequently the machining method. SEM analysis of the fracture surface showed High-Cycle Fatigue (HCF) as the failure mode. Instrumented experimentation conducted subsequently revealed a “Backward Whirl” phenomenon, initiated from the unbalance of an adjacent component and manifested at a certain range of gas generator speed, as the root cause of the HCF and consequent failures. The FRAM methodology helped in completely obviating the failures by various remedial measures, like removal of the unbalance, micro-shot peening of the shear neck, and damping of backward whirl using bearing liner bushes and enhancing stiffness by increasing the shear neck diameter to shift the backward whirl beyond the operating range of the aero-engine. The case study also demonstrates the application of FRAM for defect investigation of aero gas turbine engines.

The present work reports the measurement of planar droplet velocity field of a plain liquid jet injected into a high-speed crossflowing airstream. PIV was employed to measure the instantaneous and average droplet velocity field in the far-field region of the spray. Water was injected from a 1-mm orifice and crossflow air velocities up to 110 m/s were investigated. The approach using PIV was validated using experimental PDPA data reported in literature. The increase of droplet velocity with axial distance was clearly observed. In the region considered for analysis, the droplet velocity showed a peak in the central region suggesting contributions primarily from surface breakup.

This paper reports computational investigation carried out on T106 LP turbine linear cascade to optimize the blade performance and reduce the blade count around the LP turbine rotor by decreasing blade solidity. T106 LP turbine blade of chord 196 mm and two different blade solidities of 1.25 and 1.176 were used. Passive flow control device—Gurney flap (GF) was attached to the trailing edge of the blade. The GFs of heights 1.33% Ch and 2% Ch were used in simulations. A two-equation eddy viscosity turbulence model, shear stress transport (SST) model was considered for all the computations along with gamma–theta (γ–θ) transition model. Computations were carried out for all the cases at four different Reynolds numbers. Lift coefficient, total pressure loss coefficient, overall integrated loss coefficient, and lift coefficient to overall integrated loss coefficient ratio were used as a measure of aerodynamic performance for the cascade. From the computations, it was found that on increasing the blade spacing by keeping the GF height constant, the performance of the turbine cascade decreases. But, the performance can be improved by increasing the flap height appropriately. For the cascade configuration with increased spacing, an optimal GF height was determined.

Advanced fighter aircraft requires a gas turbine engine with high thrust to weight ratio of the order 10 and low specific fuel consumption of the order 0.7 (kg/kg-hr) to meet the high maneuverability, long range, and low life cycle cost requirements. To meet high thrust to weight ratio and low specific fuel consumption, aero gas turbine engine demands high turbine entry temperature and high turbine efficiency. In order to reduce design cycle time typically, a turbo machinery design process is carried out with the assumption that the flow is steady. However, the fluid flow in turbo machinery is highly three-dimensional and inherently unsteady due to stator–rotor interactions through wakes, potential flow, and shock interactions. In this paper, an attempt is made to analyze the unsteady flow in a transonic High-Pressure (HP) turbine which is having high blade loading and low aspect ratio, and is designed for an advanced engine. The calculations are performed by using ANSYS-CFX, which is a commercial software. This software solves three-dimensional Navier–Stokes equations. Structured grids are used in this analysis and turbulence is modeled by using k–ω SST turbulence model. Sliding interface models are used for unsteady simulation studies to analyze the flow field of the turbine stage. Numerical study shows that total-to-total efficiency of the HP turbine stage decreases by 0.4% due to unsteadiness as compared to steady state.

The substantial adverse pressure gradient experienced by a turbulent boundary layer while approaching an endwall-mounted twisted turbine blade and caused the impending flow to separate three-dimensionally to form a dynamically active horseshoe vortex (HSV) system in the junction of the turbine blade with endwall. The large eddy simulations (LES) of the flow past a twisted turbine blade mounted on a curved endwall with periodic boundary condition in pitchwise direction is carried out for Re = 50000 to methodically investigate the HSV dynamics. The significant variations with Re in terms of mean flow quantities, heat transfer distribution, and coherent dynamics of turbulent HSV are shown in computed results. The HSV system consists of a multiple number of necklace-type vortices that are shed periodically at maximal frequencies. For high Re, we show that outburst of wall govern the instantaneous flow field, averaged vorticity affiliate with the growth of hairpin vortices that enclose around and dislocate the primary HSV. The time-mean endwall heat transfer is prevailed by two bands of high heat transfer which encircle the leading edge of the blade. The band of maximal heat transfer, occurs in the corner region of the juncture, while the secondary high heat transfer band (thin as compare to primary) develops upstream of primary band, in between primary and secondary bands a relatively low heat transfer region is identified.

In the present work, a single-stage axial turbine is designed using three design approaches for the same specifications. The three different designs were obtained using different classical vortex distributions, viz., Free Vortex (FVD), Constant Nozzle Angle (CNA), and Constant Specific Mass Flow (CSM). Kacker–Okapuu model is used for the estimation of pressure losses. To maintain consistency, the design is carried out for the same flow path and turbine stage parameters, viz., the stage loading, flow coefficient, and mean reaction. The hub-tip radius ratio is 0.72 for all the designs. The design point performance and flow analysis are carried using a commercial CFD solver. A comparative study of the performance obtained using the three different approaches is carried out. It is observed that for a turbine with a hub to tip ratio of around 0.7, the choice of vortex distribution does not yield any notable difference in efficiency output at design point.

The operating gas temperatures of advanced military aero engines are continuously increasing to achieve a higher specific thrust. The gas temperature exposed by hot end components is beyond material allowable temperature limits. It is essential to cool these components to lower the metal temperatures to meet the designed life. Effusion cooling technique effectively brings down the hot components’ metal temperature. The present study is focused on the numerical prediction of cooling effectiveness from effusion cooling holes. Conjugate heat transfer analysis is carried out using ANSYS Fluent ver.14.5 to estimate the film cooling effectiveness. The analysis is validated with experimental film cooling effectiveness data. K–ω SST turbulence model has given good agreement. To improve the film cooling effectiveness, further studies have been carried out by varying the geometrical parameters such as film hole inclination, porosity and thermal conductivity of the effusion plate. Three cases with different cooling configurations have been investigated.

Considerable efforts have been made by aero-engine manufacturers to understand, detect, attenuate, if not eliminate, high-frequency transverse screech combustion instabilities due to their destructive nature. Combustion dynamic stability problems arise in turbofan afterburners when fluctuations in the combustion energy release rates are coupled with the afterburner duct acoustics. Anti-screech liners have been used to attenuate the consequential pressure oscillations and mitigate the harmful effects of the transverse screech combustion instabilities. A versatile experimental test facility, using a single V-gutter flame holder as a driver to generate predetermined screech frequencies of interest of 1250 and 2000 Hz was used to conduct comprehensive experimental studies. Anti-screech liners of variable effectiveness with porosities of 3.5 and 10.0% were used for this investigation. The attenuation effectiveness of these liners was investigated under simulated test conditions specifically for the 2000-Hz screech frequency. For these comprehensive experimental investigations, a novel methodology to generate the predetermined screech frequency in a model afterburner test rig had been evolved. These anti-screech liners with variable effectiveness were not found to be effective in completely suppressing screech over the entire spectrum of operating conditions, probably because the complex modes present in the screech combustion instabilities, may have not been strictly transverse and the pressure amplitude may have been too large. Mitigation efforts need to be attempted at the source of generation of the screech combustion instabilities.

Military aircraft employ aero-gas turbines fitted with afterburners to meet the requirements of rapid increase in thrust for flight operations that involve combat maneuvers. The airflow rate through an aero-gas turbine remains unchanged even after an afterburner is invoked to ensure that there is no disruption in the turbomachinery operating characteristics. Triggering of flow instabilities leading, in turn, to combustor instabilities could occur due to the incorrect Rayleigh heat addition in the constant area afterburner. In normal operation, the propelling nozzle should be correctly opened up to pass the increased volumetric flow rate of the heated air due to heat addition in the afterburner. Any mismatch could result in violent flow instabilities including possible fan stall in a turbofan. The processes of triggering instabilities by sub/supercritical Rayleigh heat addition have been characterized. The gas dynamic equivalence of secondary mass addition to heat addition has been analyzed and experimentally validated in a model afterburner combustion test rig. Consequently, sudden sub/supercritical heat addition in an afterburner with its corresponding propelling nozzle closure has been studied and characterized for the equivalent mass addition in a separate model afterburner simulation test rig.

In-flight flame out of a gas turbine engine is one of the worst nightmares of the aviators. Relight of the engine is the most benign action possible to restore the safe flight. In such circumstances, the starting process is a vital issue that needs careful study. The starting phase of the gas turbine engine is a highly complex and transient phenomenon. It involves the core aero-thermal portion of the engine and systems such as ignition system, fuel system, and control system and in some extreme cases of relighting assistance from starter also required. As far as the turbomachinery is concerned, the starting phase is an off-design condition, except for the starter and ignition system. From the flight safety point of view, the importance of reliable engine start, especially the in-flight relight of aircraft engine needs no explanation and assurance about midair restart capability need to be proven beyond doubt. The methodology demonstrated in this paper is a novel tool that can be used by maintenance engineers to identify the impending failures of the starting system. This paper highlights the outcome of the work carried out for a detailed analysis of the starting process via the starting trials of turboprop engines. Based on the detailed study of the starting process carried out during initial aircraft integration of the turboprop engine, the authors could identify the faults in the starting system of turboprop engine in the ground which has an effect on safe and reliable air starts. In-flight relights data and ground starts data have been analyzed as a part of this study to identify the novelties in the starting process so as to detect the faults.

Small gas turbine engines are increasingly used in cruise missile applications. In the design stage of these engines, aero-thermodynamic models are used to evaluate the expected performance of the engine for a given set of component characteristics. The throttle characteristics and altitude-Mach number characteristics of the engine are iteratively analyzed using this model. This paper describes the use of such a model to show the expected performance of a set of design choices at different altitudes and Mach numbers. These small engines are operated at maximum possible Turbine Inlet Temperatures (TIT) for maximum thrust. Theoretical relations that give the slope of the operating line for constant turbine inlet temperature operation is derived. Using these expressions, the reduction of stability at high altitude and low Mach number is shown.

Rotor blade vibration in turbomachinery has been a major cause of failure due to HCF, often resulting in catastrophic damage. The primary aeromechanical design concerns are blade flutter and forced vibration that need to be quantified. The severity of blade vibratory response is almost impossible to predict using theoretical tools as it depends on the strength of excitation. Hence in order to evaluate the HCF characteristics of rotating blades, aero industry depends on measurements for actual vibratory response during engine tests. Various methods are used for measurement of rotor blade vibration. Conventionally strain gauges are extensively used for characterizing vibratory signatures of rotating blades. However, the strain gauges have their own limitations posed by operating temperatures and high-end technology is required to transmit signal from rotating components. Hence only a few blades in a rotor can be instrumented resulting in limited data capture. This paper presents a non-contact type of measurement technique using blade tip timing to capture vibratory signatures of all the blades of the rotor stage. This method is used to characterize monitor rotor blade vibrations of Low-Pressure Compressor and Low-Pressure Turbine of a developmental gas turbine engine. It has provided valuable data with respect to incipient damages, preventing catastrophic failure.

Developmental testing of gas turbine propulsion systems involves iterative experimental studies right from First Engine to Test (FETT) till production release. During initial design validation tests, engines require dedicated instrumentation for carrying out online monitoring with a capability to detect and isolate impending failures which may lead to structural damage. Instrumentation may be optimized further over the developmental lifecycle to determine the general engine health and life consumption of critical parts. Instrumentation generally focuses on monitoring of structural and aerodynamic behavior of engine subsystems. It is challenging to arrive at optimum instrumentation and methodologies of measurement with respect to engine performance and structural health monitoring. Structural health monitoring of rotating engine components poses challenges in acquiring high bandwidth data through either contact or non-contact sensing techniques and further data processing. Special instrumentation systems used for measuring various parameters from rotating parts as a part of health monitoring include slip rings, rotating telemetry, and non-intrusive strain measurement systems. Instrumentation of other engine parameters includes temperature, pressure, rotational speed, casing vibration, control actuator positions, flow rate, clearance between stationary and rotating parts and lubrication oil quality. Gas turbine engines are a complex assembly of rotating and stationary parts which operate at extreme temperatures which limits the operational capability of sensors. It is a challenging task to error budget complete measurement chains and arrive at uncertainties. Various aspects discussed in this paper are complex and inter disciplinary in nature. This paper provides a bird’s eye view of the challenges associated with measurement systems during developmental stage of a gas turbine engine.

This paper presents a summary of research toward extending the flight duration of fixed-wing unmanned aerial vehicles at ADE. A historical context to extended flight is provided and particular attention is paid to research in establishing the best operating profile for the Reciprocating Piston engine to meet the extended endurance target. With the limitation of the fuel capacity in the UAV, it is imperative to identify and operate the UAV at the lowest fuel consumption regime without compromising the mission objectives. Usage of variable pitch propeller along with operation of mixture control of the engine has resulted in the extension of endurance of up to 12 h. Autonomous flying presents a unique set of challenges whereby the flight computer of the aircraft must command and control the engine to be operated at the best operating regime. The basic mechanisms of variable pitch propeller operation are examined. During the present work, automation tables were evolved to aid the flight control computer.

Multi-hole pressure probes have a wide range of applications in terms of measurements in special applications like turbomachines flow field, wind tunnel experimentation applications, etc. The primary measurement variables of interest are static pressure, dynamic pressure, three velocity components, flow direction, etc., which we can measure using a multi-hole pressure probes with higher accuracy. The applications of such probes in measurement fields demands for higher accuracy of calibration, especially in terms of flow angularity and precision. The manual calibration of the multi-hole probe is a difficult and time-consuming task in an acceptable range of pitch, and yaw angles. The present paper discusses the low-speed multi-hole pressure probe calibration facility developed at IIT Kharagpur using a two-axis angular traverse mechanism. Special algorithms using LabVIEW were developed for automatic traverse both for yaw and pitch direction and also for pressure data acquisition using pressure scanner, to achieve precious measurement with short time spend with angular precision in terms of 0.5°. This facility is capable of calibrating multi-hole pressure probes for different Reynolds numbers and angular ranges. The paper also discusses the initial calibration of the tunnel using total pressure probe rakes. A four-hole probe was calibrated using non-nulling calibration technique using the aforementioned test facility. It was calibrated for yaw and pitch angle range of ±30° and ±70°, respectively. In this study, we are discussing three calibration methods, namely, the conventional method, the two-zone method and three-zone method to understand their behavior in terms of calibration coefficients, operable angular range, and uncertainty. The range for the pitch angle of the conventional method was observed to be $$\pm 35^\circ$$ . However, an extended range for pitch angle up to $$\pm 60^\circ$$ for two-zone method and three-zone method was observed. The uncertainty analyses of the results have been performed to study the sensitivity of the probes at prescribed angular ranges.

An experimental study is conducted to investigate the performance of a non-uniform area resonator-based acoustic pump for air. Higher peak pressure amplitudes can be obtained by employing a non-uniform resonator. The pressure obtained has an oscillatory nature about the ambient pressure, with the frequency of operation equal to 933 Hz. The extraction of useful energy from the resonator is achieved by no-moving-part valves. In the present study, four different configurations of no-moving-part valves have been used for rectification. A series combination of three 15 mm length flow individual diode elements showed expected results with the diode having smaller diameter exhibiting better rectification. However, single diodes with length of 45 mm exhibited rectification in the opposite direction. The performance of the acoustic pump is evaluated based on its tank filling characteristics for a given driving power and a particular valve configuration. Maximum tank pressures of the order of 1000 Pa are obtained at maximum flow rates of the order of 10 slpm for the available range of driving powers.

This article discusses about schlieren methods which can draw hidden details in high-speed flow visualization. Mainly, two schlieren techniques are used, namely, inclined schlieren and focusing schlieren. These two techniques have been applied to visualize normal-shock–boundary-layer interaction in ducts. The paper explained the possibility to measure normal shock oscillation in spanwise/transverse direction. Identification of shock location across the duct span is also illustrated, using focusing schlieren. Further, the possibilities and limitations of focusing schlieren in high-speed flow visualization have been discussed.

Large field flow visualization of the unsteady combusting flow inside a hydrogen-fueled valveless pulsejet engine has been successfully demonstrated by using a retro-reflective screen based high-speed shadowgraph technique. A rectangular cross-sectional valveless pulsejet engine with optical access has been designed, fabricated, and successfully used for demonstrating the effective use of a retro-reflective shadowgraph technique for large and spatially wide flow fields of interest. The aspect ratio of the engine considered for the study was 13.5 and the technique helped to understand the acoustically coupled combusting flow structures from the intake to the tailpipe of the pulsejet engine. A Photron FASTCAM SA4 camera was used in this study. High-speed shadowgraph videos were captured with a frame rate of 13500 frames per second with a resolution of 1024 × 272 pixels.

Screech combustion instabilities continue to be one of the most detrimental and, in fact, fatal issues for the development of a high-performance afterburner. It is essential to study the underlying flow-physics related to the crucial thermo-acoustic coupling to acquire a predictive capability and to evolve a methodology for the attenuation of screech. A versatile test facility using a single V-gutter flame holder was used to generate the predetermined screech frequency of 2000 Hz in a controlled and sustained manner. The test facility had the capability to run the afterburner model under simulated inlet conditions of pressure and temperatures. The critical zone of flame stabilization near the V-gutter flame holder had complete optical access with quartz glass windows to study the vortex shedding phenomena during the afterburner operation. The critical operating parameters of the test rig were measured using a high-speed NI-based data acquisition system. Flow visualization studies using a high-speed shadowgraph technique was effectively used to understand the onset of screech. A FASTCAM SA-4 Photron high-speed camera was used in this experimental investigation. It was found that the cause for the onset of screech was the vortex shedding frequency from the flame holder locking on to the duct transverse acoustic resonant mode frequency.

In aircraft engines operating at high Mach numbers, it is exigent to reduce the magnitude of flow speed from supersonic to subsonic before entering the burner, to accomplish a proficient ignition. It is accomplished by a progression of oblique shocks as well as a normal shock wave in supersonic intakes. However, the advantage of shock enabled compression in intakes does not come independent but with colossal losses due to shock-boundary layer interactions, which includes intake unstart and an abrupt thickening/separation of the boundary layer. Controlling these interactions by boundary layer control using micro-vortex generators (MVGs) has gained prominence. In the present study, an attempt is made to control the interactions at the shock impact point in a Mach 2.2 mixed compression intake. Two types of MVGs; a conventional configuration and an innovative ramped-vane configuration were experimentally investigated by varying the MVG heights as 600, 400, and 200 μm. Also, the near-field effects of MVGs are quantified in terms of static pressure variation in the internal flow. It is found that the innovative MVGs of height 200 μm leads to favorable pressure drop in both the upstream and downstream region, due to enhanced flow mixing near the boundary layer.

Experiments were performed to study the effect of cavity rear wall modification at different L/D ratios of an open cavity. All the experiments were carried out at Mach 2.0 and corresponding Reynold’s number of 5.5 × 105 based on cavity depth. Unsteady pressures were measured at several locations to obtain the fluctuating flow field behavior and the pressure spectrum. Results indicate that cavity rear geometry modification has its influence toward the instabilities and subsequent flow oscillations inside the cavity at all L/D’s.

Scramjet engine is the most promising air breathing propulsion system in the hypersonic flight regime. Combustion in a scramjet engine, however, is difficult to achieve due to flow residence times being comparable to chemical times. For our study, 3-D RANS CFD analysis of a strut flame holder of a scramjet engine combustor was carried out. The first part of the analysis focuses on the validation of CFD gas dynamics results with earlier experimental data in literature. The air enters the combustor at Mach 2 with a stagnation temperature of 612 K and a stagnation pressure of 7.82 atm. Various flow properties like velocity profile, wall static pressure and density gradient were in good agreement with experimental results. The next part of the analysis involves a strut mounted across the walls of the combustor with hydrogen fuel injector jets forming an aerodynamic ramp on both the upper and lower surfaces. The location of the injectors relative to the strut base and dynamic pressure ratio were varied and the observations presented.

To sustain the arc in the arcjet, the swirl is thought to be necessary. This paper conducts a numerical study on various geometry and input conditions. The effect of swirl and wall temperature of a nozzle on the plasma flow, arc behavior, and performance parameter of low power arcjet has been studied. The physics behind the working of low power arcjet and behavior of swirl velocity has been plotted. Various aspect of swirl, which stabilizes the arc, is also discussed.

Hypersonic airbreathing propulsion has applications in defence and space accessibility. The objective of this study is to analytically model a Dual-Mode Scramjet with fuel flow rate as the controlling parameter at a design Mach number of 6 with Kerosene as fuel. The components of the scramjet were modeled individually and integrated through a one-dimensional analysis. The location of the fuel injectors determines the axial distribution of stagnation temperature in the combustor. This is necessary to evaluate the location of thermal choking. The combustor-isolator model can accommodate supersonic and subsonic combustion with the intermediate mode involving a Normal shock in the combustor. An algorithm was developed to determine the axial location of this Normal shock. On mapping the fuel flow rate to the thrust and integrating it with the Drag model, the flight profile of the vehicle can be simulated. Finally, the feasibility of controlling the trajectory (Mach number and Altitude) of a DMR-powered vehicle with the fuel flow rate as the input parameter is assessed analytically.

In the present study, attempts were made to study the effect of activated charcoal (AC) on performance of hybrid rocket motor with polyvinyl chloride–di-octyl phthalate (PVC–DOP) as solid fuel. Activated charcoal at 1, 2 and 5% mass fractions was added to PVC–DOP. The regression rate and combustion efficiency of the PVC–DOP/AC fuel combinations were determined using laboratory-scale hybrid rocket motor with gaseous oxygen as oxidizer. The results showed that among the fuel combinations studied, PVC–DOP with 1% AC exhibited the highest improvement in regression rate as well as combustion efficiency. The excessive char formation was observed as the mass fraction of AC increases and this causes the reduction in performance of hybrid rocket. Additionally, the mechanical properties were studied and TGA analysis was also conducted for the PVC–DOP/AC fuel samples.

Gas turbine engine carcass is one of the major structures that supports/houses the main engine rotors, stator blade rings, interface features to mount the engine onto the airframe and accessories, viz., gearbox, electronic control units, pipelines, actuators, and so on. All the reaction forces generated by the engine are finally grounded to the airframe through this engine carcass. Specific to aero engines, designers make every effort to reduce the weight of this engine carcass so as to maximize thrust to weight ratio of the engine while maintaining the stiffness requirement. The present paper deals with an integrated approach for health assessment of the engine carcass. The theoretical and experimental studies undertaken for the engine carcass and rotor for enhancing the diagnostic reliability have been put forward. The requisite instrumentation and signal processing to assess the health have also been included. The decision-making for furthering of the test and remedial measures are also brought out.

The structural integrity evaluation of the aero structures and components is of paramount importance not just to validate the design and manufacturing processes, but it is also mandatory to prove the satisfactory compliance of design objectives laid down with respect to the performance and functioning. The paper depicts the methods employed for simulating the bird-strikes on the aero structures (of both military and civilian aircrafts) in the state-of-the-art test facility developed for conducting such tests along with some test cases and conceptualizing a new technique for the high (linear) speed measurement of the bird.

Squeeze film dampers play a vital role in absorbing vibration energy in a rotor bearing system. The damper under study has an elastic ring with pedestals between bearing and stator dividing oil cavity into small oil pockets. This arrangement is different from conventional squeeze film damper where a single annular oil film is formed. This provides the required support stiffness as well as damping to the rotor. These types of dampers are called elastic ring squeeze film dampers (ERSFD) which are mainly used in high-speed small gas turbine engines by virtue of its compact design. There are very few literatures available to evaluate the damping offered by these SFDs. The main objective of this work is to determine the damping offered by ERSFD experimentally. For this study the rotor is designed to simulate the dynamics of a typical gas turbine engine. The rotor has to cross two rigid critical speeds within 18,000 rpm. The rotor response is measured under undamped (UND-without oil supply) and damped (D-with oil supply) conditions to evaluate the damper performance. The performance data is generated at three different oil temperatures (40, 70 and 100 ℃) under unbalanced load ranging from 2 to 8 g at 58.5 mm radius (UBR). This experimentation and performance analysis shows enhanced damping at critical speeds leading to the reduction in rotor vibrations after introducing ERSFD. The experimental data is further processed to calculate the amount of damping offered by ERSFD using rotor dynamic relations.

Low stiffness bearings are useful to reduce the force transmitted from the vibrating rotor to the surrounding support structure. However, having low stiffness requires us to cross the low rigid body critical frequency while accelerating to operating rpm. In this work the stiffness of the bearing is changed online during operation by using an active magnetic bearing instead of a conventional constant stiffness rolling element bearing. This methodology is shown for a rigid rotor using both simulation and experimental techniques. During acceleration phase, a high stiffness is maintained, which gives us high critical frequency. After acceleration to operating rpm, the stiffness of the bearing is reduced at run time so that the bearing again becomes a soft support. In this work, a thrust magnetic bearing of variable stiffness is used to show that by changing the stiffness at run time, we can avoid crossing the rigid body critical frequency and hence reduce the amplitude of resonant vibrations.

The objective of this study is to investigate the suitability of Aramid fiber polymer composite as reinforcement for blade containment in a gas turbine system. The preliminary experimental study is focused on optimization of Aramid fiber epoxy laminate composite in terms of number of layers and fiber orientation with respect to energy-absorbing capability. Acoustic energy attenuation study has been performed along different directions of laminates to gauge the effect of fiber orientation and number of layers on stress wave attenuation. NDE measurements and standard destructive tests were performed on sets of samples with varied number of layers and orientation of Aramid fibers. Energy-absorption characteristics of the Aramid fiber epoxy composite laminates along with these variables were investigated. This experimental study has yielded some interesting and encouraging results.