Design and Development of Aerospace Vehicles and Propulsion Systems

The present experimental investigation reports the effect of chevron on flow characteristics and associated acoustic characteristics at different jet exit Mach numbers in high subsonic compressible jet. Flow characteristics are investigated with mean pressure measurement using miniature pitot tube in the flow, and acoustic characteristics are investigated using fluctuating pressure measurement with array of four microphones in the far field of jet. Chevron is used as a flow control device on the lip of the nozzle. Chevron converts axisymmetric jet development into corrugated shear layer, closer to the nozzle exit. This effect diminishes as jet grows downstream away from the nozzle exit. Corrugation of jet shear layer closer to the nozzle exit increases with increase in Mach number. Compressibility effect with change in Mach number is seen from potential core length and jet growth rate for base nozzle; however, chevron is found to reduce the compressibility effect with change in jet exit Mach number due to enhancement in mixing. Chevron reduces far-field overall sound pressure level at shallow polar angle (30°) by about to 2 dB at all the Mach number; however, increase in noise level at higher frequency observed at higher polar angle is mainly due to high-frequency noise sources produced from chevron petals. Noise level at higher polar angles and higher frequencies increases with increase in Mach number.

Shock wave-boundary layer interactions take place in many vehicle configurations of practical importance such as wing–body junctures, deflected control surfaces, high-speed inlets and forward-facing steps. The associated flowfield becomes complex when the interaction causes flow to separate. A parametric study is carried out for supersonic flow past a compression–decompression ramp (CDR), to determine the ramp inclination angle that minimizes the adverse effect of separated flow. The inclination angle below which weak interaction occurs is obtained from the numerical simulations.

In the course of an aircraft development program of LTA, a number of engine-related ground and flight tests have to be carried out to determine the losses particular to the engine installation. These installation losses must be determined in order to create the final aircraft power setting charts. During the course of configuration development, design analysis is a must to ensure that the proposed design will perform the intended operation within the estimated losses, this confidence should be assessed either in a bench tests or in CFD analysis. The numerical study of engine air intake performance is carried out using RANS-based $$k - \omega$$ k - ω SST using ANSYS FLUENT software. The study is carried out for minimum climb condition, as it is critical operation of any aircraft mission. The CFD results were compared with the flight test data for validation. The instrumentation that is necessary to measure installation losses and engine performance has been installed. Computational results agree well with the experimental results and were found satisfactory. The CFD results of air intake show that the total pressure is extensively recovered at the engine intake plenum (compressor inlet screen) region, and the inlet pressure loss and ram air recovery were within the acceptable limit as recommended by engine OEM.

Aero-elasticity of a wing is the study of airflow around an elastic wing and their interaction. A two-way Fluid Structure Interaction (FSI) method has been advantageous in the study of aero-elastic behavior of lifting surfaces. Preliminary design of an UAV requires detailed understanding of aerodynamic and structural operational boundaries; to arrive at an optimal wing. High aspect ratio wings are being employed in Medium Altitude Long Endurance (MALE) and High-Altitude Long Endurance (HALE) UAVs. High aspect ratio wings are usually designed using composite materials because of their specific stiffness and strength customization ability. To this end, for designing a high aspect ratio wing, which is aero-elastically compliant and performs well at all flight conditions—understanding the aero-elastic behavior is necessary. Preliminary designs as well as detailed design incorporate reduced ordered models, however to evaluate the efficiency of reduced ordered models—a full scale 3D numerical model analysis or wind tunnel tests are required. Here aim is to perform and establish a baseline study of two-way coupled static aero-elastic analysis based on the existing experimental results of a composite flat plate wing and a large/high-aspect ratio wing, which in turn will work as a benchmark for flutter analysis and reduced ordered models to be developed. ANSYS package has been used to demonstrate the methodology to predict the onset of divergence, a static aero-elasticity phenomenon. The numerical simulation results from ANSYS are validated with existing experimental results available in the literature. Numerical simulation is carried out for a composite wing, and then the methodology is extended to high aspect ratio wing. The results obtained using bidirectional-coupled simulation studies are found to be in agreement with the results available in the literature.

The present numerical study deals with the steady and unsteady aerodynamics of an airfoil in a cascade with oscillating neighboring blades. The motivation for the study arises from aeroelastic studies of turbomachinery blades, where the unsteady forces acting on a blade due to different sources is examined. In order to identify the root cause of the various phenomena affecting the aeroelastic stability, it is necessary to observe the effect of only one source of disturbance and isolate the others. The cascade comprises of five blades with zero stagger and low incidence. The two adjacent blades to central one are oscillated with a fixed frequency and phase difference, with the rest of the blades remaining stationary, covering a range of frequencies and phase difference angles. The primary objective is to look at the variation of global parameters on central blade with reduced frequency. The moment and drag are estimated numerically and compared with experimental results which shows good agreement. The hysteresis loops of lift and moment coefficient with angular displacement are used to understand the effect of reduced frequency. A laminar separation bubble is observed to be formed during part of the oscillation cycle and its size is related to the unsteady forces on the blade. The behavior is also a function of reduced frequency to some extent. Such an understanding of the effect of oscillating blades in a blade row is essential for the modeling of the aerodynamic forces in an aeroelastic problem.

Modern civil aircraft engines are known for their high bypass ratio fans that are powered by many low-pressure turbine (LPT) stages (Mahallati et al. in J Turbomach 135/011010, [1]). LPT can contribute as much as 30% of the weight of an aero engine (Sondergaard et al. in Toward the expansion of low-pressure turbine airfoil design space, [2]; Curtis et al. in J Turbomach 119(3), [3]). Aerofoils of modern LPT blades are subjected to increasingly stronger pressure gradients as designers require higher blade loading in an effort to reduce weight and costly number of LPT blades of an engine, which leads to better reliability and maintainability. This decrease in number of LPT blades results in increase of thrust/weight ratio, thus reducing the fuel consumption. But highly loaded aerofoils can reduce aerodynamic performance, and its influences can be seen more in unsteady environment. Hence, reduction of losses in unsteady environment, improves the turbine performance which is a challenging task. Therefore, there are persistent efforts towards the generation of “high-lift” blade profiles. As a result, industry and research communities are motivated for further deep research efforts in LPT aerodynamics (Sarkar in J Turbomach ASME 131, [4]). In this paper, the effect of the incoming wakes shed from the upstream HPT blade on the downstream highly loaded transonic LPT vane are studied to better understand the LPT flow physics. In this paper, wakes shed from upstream HPT blade are simulated by cylinders of three different radius which are of the order of actual HPT rotor blade trailing edge radius ~0.8 mm, so that only the influence of wakes on vane losses can be studied. From these three cases, it is observed that incoming wakes does not always lead to increase in vane loss coefficient. For some cases, it is observed that the loss coefficient is 12.64% lower than the vane without any incoming wakes.

Control jets injected into a supersonic flow cause a significant region of separated flow over the parent vehicle in the vicinity of injection. This altered external flow contributes to an additional force on the vehicle in addition to the jet thrust thereby effecting its performance. In this study, numerical simulations were carried out to estimate the effect of injection location and body attitude on the overall performance of a Reaction Control System jet in producing a control force. Commercial CFD software was used to solve the 3D RANS equations using the SST-k $$\omega $$ ω turbulence model. It is seen that favourable interaction results from injection at rearward locations and positive angles of attack whereas injection from forward locations and negative angles of attack results in an adverse interaction.

Current paper deliberated the impact of gap-to-chord ratio (g/c) variation on grid fin subsonic flow characteristics through chord modifications while retaining the same gap and aspect ratio to explicitly decipher the role of chord on grid fin aerodynamics through numerical analysis. Solver validation is followed by comprehensive examination of the pertinent aerodynamic coefficient results associated with different grid fin chord variants. The study establishes enhanced maximum lift coefficient at the expense of reduced aerodynamic efficiency in the most operable angle of attack region for higher g/c (lower chord). Efficiency reduction is attributed to increase in drag for higher g/c emanating due to increased pressure drag applicable for a blunt compact geometry. Stall angle undergoes minimal deviation for different g/c when chord is the varying parameter. However, deviations associated with grid fin aerodynamic efficiency for varying chord were found to be appreciably significant. The study categorically deduces the impact of chord in the g/c parameter, and hence can be helpful for grid fin designers while selecting the optimum chord value for enhanced aerodynamic efficiency and lift requirements. This study in conjunction with analysis carried out for gap variation can help achieve an efficient grid fin design with respect to delayed stall angle and increased aerodynamic efficiency.

The flow over a flat plate airfoil with 5-to-1 elliptical leading and trailing edge at Re = 80,000 and for different angles of attack (0°–15°) is numerically investigated by solving the Reynolds-averaged Navier–Stokes equations. The $$k-\omega$$ k - ω shear stress transport equation, $$\gamma -{Re}_{\theta }$$ γ - Re θ turbulent transition model is used to address the effect of laminar-turbulent transition. The present computed aerodynamic forces are compared with the available experimental data for validation. Large flow separation and a single recirculation zone is found at higher angles of attack. The study is extended to investigate the laminar separation bubble effect on a three-dimensional Zimmerman wing planform for Re = 50,000; 100,000 and 150,000 at different angles of attack. The present results agree well with the available experimental and computational data. The influence of propeller on the aerodynamic performance of a Zimmerman wing planform is investigated. The results show that the wing with propeller configuration has lower $${C}_{D}$$ C D values compared to wing alone case. The results presented in this paper show the importance of modelling the propeller slipstream effects on the aerodynamic characteristics of low aspect ratio wing.

The present study focuses on the design and performance evaluation of a rectangular mixed compression intake for ducted rocket ramjet applications. For the theoretical design stage, a 1D optimization criterion has been used to fix the ramp angles for theoretical maximum total pressure recovery (TPR), at on-design M 2.9 (shock-on-lip). Two different ramp designs (two and three ramps) have been considered for optimization, to find their effect on the overall performance. The throat height has been fixed using the practical self-starting contraction ratio (CR), (i) for the on-design Mach number of 2.9 and (ii) for the low off-design Mach number of 2. The throat length used is about six times the throat height for the required supercritical margin as well as to contain the shock train and the subsonic diffuser divergence angle is about 6°. The viscous flow field has been obtained by solving Favre averaged Navier-Stokes (FANS) equations with two-equation SST k-ω model. The analysis shows that, for the two-ramp and three-ramp design with self-starting contraction ratio at M 2.9, the on-design critical TPR is 0.645 and 0.67, and the critical mass flow ratio (MFR) is 1 and 0.99, respectively. The performance at low off-design M 2 shows, the MFR of both designs reduces to 0.51 and 0.47, respectively, and improves to a value of 0.6 for the two-ramp configuration with self-starting CR at M 2. This indicates that the two-ramp design has a better low off-design Mach number performance.

Staging of a launch vehicle is essential for achieving the required orbital velocity and altitude. For the launch vehicle studied here, the boosters are detached from the ongoing stage (core) by means of jettison motors. The jet pressure ratio is high and the plume from the jettisoning motors expand and propagate upto the heat-shield region. Thereby aerodynamics of the complete vehicle alters. The present study is also focused on the effect of altitude variation on the aerodynamics of the launch vehicle during hot separation. Two different trajectories (or altitudes) are studied and the effect on the aerodynamic coefficients is quantified. The paper presents in detail about the simulated conditions, the variation in aerodynamic coefficients and the complex flow-field involved.

The normal shock train oscillations in a rectangular duct were studied experimentally. The studies have been classified into three groups. The studies in group I showed normal shock train oscillations and its structures at different duct locations. Group II studied about the shock oscillation when it stands on a hybrid micro vortex generator, while group III studied moving normal shock train. The shock structures have been visualized using high-speed schlieren imaging, and its oscillations have been quantified with time resolved image processing. It has been showed that the tertiary shock can oscillate at larger amplitude than secondary shock. Moreover, the injection assisted MVGs reduces shock oscillation even at its off-design condition.

A special side wall mounted test set up rig was designed and realized for hinge moment characterization of an all movable control surface. Experiments were conducted on a 1:2 scale down model of the control surface. In the experiment, two different configurations (A & B) were tested. In configuration A, the control surface is mounted on its actuator cover whereas in the configuration B, the control surface without actuator cover is tested. In case of configuration B, a boundary layer bypass plate was used between tunnel test section side wall and the control surface. The test Mach numbers are 2.0, 2.5, 3.0, and 3.5. The control surface was deflected from −10° to 25° during the test using electrical stepper motor. Integral strain gauge force balance has been used to measure the forces and moment on the control surface. Test data have been analyzed. Normal force co-efficient, hinge moment co-efficient, and variation of center of pressure have been presented in this paper. Hinge moment is significantly influenced by the presence of actuator cover, and it is found higher compared to configuration B. Center of pressure is ahead of the hinge line toward the leading edge of the control surface for both the configurations. The hinge moment is unstable for both the configurations. However, it is more unstable in case of configuration A.

The present paper reports findings of the 3D computational studies of the effect of frontal gusty shear flow on the force patterns of a flapping wing. A rigid wing with semi-elliptical wing planform with asymmetric 1 DoF flapping kinematics was exposed to a gusty shear flow. The shear gradient of the flow was varied from −10 to +10 in steps of 5. Computation studies were carried out for Re = 150 which lies in the typical flight Reynolds number range of natural flyers like a fruit fly and anthropogenic flyers like a Pico Aerial Vehicle. 3D, unsteady, laminar, and incompressible Navier-Stokes equations were solved using finite volume formulation based commercial code ANSYS Fluent. Wing kinematics and gusty inflow conditions were modelled into the solver by User Defined Functions (UDFs). Wing motion was simulated using the dynamic meshing technique. The effect due to variation in the frontal inflow condition was studied quantitatively and qualitatively. Comparisons of the instantaneous and gust cycle averaged forces and moment coefficients about the wing root mid chord point and 3D phase space projections of the forces and moment coefficients was carried out. Qualitative studies were carried out by comparing the static pressure over both the surfaces of the wing and the vortex patterns near the flapping wing using λ2–criteria. It was observed from these studies that negative shear gradient resulted in a rise in the vertical force and moment and a minor reduction in the horizontal force. Positive shear gradient resulted in a minor rise in horizontal force and a significant reduction in vertical force and moment.

Aerodynamic coefficients for a launch vehicle are estimated using wind tunnel tests as well as CFD simulations. Estimation of coefficients at liftoff is often carried out using low speed wind tunnel tests. Presence of tunnel walls affect aerodynamic coefficients due to the blockage introduced by the model to the flow. The effect of blockage is investigated in the current study through CFD simulations with and without wind tunnel walls on a launch vehicle configuration in the presence of launch pad and umbilical tower (UT). It is seen that the aerodynamic characteristics are significantly influenced by the presence of wind tunnel walls. Even though the model blockage is 4%, variation of 8.5% is seen in total force coefficients.

An experimental study has been conducted in a Mach 2.5 flow to understand the shock-wave boundary layer interaction (SWBLI) over a forward-facing step (FFS) of step height equals twice the oncoming boundary layer thickness. Particle Image Velocimetry (PIV) is the primary methodology used to study the unsteady shock oscillation phenomena. Prior studies show shock oscillates one or two orders smaller than the characteristics frequency of the incoming boundary layer. Particle Image Velocimetry has done both in cross-stream plane and spanwise plane. Instantaneous PIV images in the cross-stream plane shows shock foot oscillates over a distance of approximately one step height. From PIV study in the cross-stream plane, it is observed that the incoming boundary layer velocity fluctuations (upstream parameter) is corelated to the shock foot oscillation and the separation bubble (downstream parameter) is correlated to the shock motion for the forward-facing step of step height equals twice the incoming boundary thickness.

RANS computations are performed for M∞ = 8.2, Re∞/m = 9.35 × 106 flow past a flat surface with protuberance of size of the same order as that of the boundary layer. The local intensification of heat transfer rates in the vicinity of the protuberance is considered as the main focus of the study. The effect of separated and unseparated boundary layer is considered by varying the protuberance angle. The results are compared with available experimental data. The maximum heat transfer rate, a major design parameter, is found to be fairly predicted by computations. However, accurate prediction of parameters such as upstream separation length and heat flux distribution requires further research and calibration of available turbulence models.

Aerodynamic shape optimization is carried out for (a) maximizing lift coefficient, Cl and (b) maximizing endurance factor (Cl3/2/Cd) of an airfoil at a transitional low Reynolds number (Re = 5 × 104) using surrogate-based optimization technique. Bezier curves are used to parameterize the airfoil. Latin hypercube sampling (LHS) technique is used to sample the initial set of samples. The aerodynamic response is estimated using low- and high-fidelity datasets. A co-kriging-based multi-fidelity surrogate is built using these datasets, and the response surface is used in the optimization procedure for the search of optima. Expected improvement strategy is used to update the surrogate at every iteration, and optimization is terminated once the convergence criterion is met. The objective function improvement for maximizing Cl and maximizing (Cl3/2/Cd) is 18.08% and 38.70%, respectively.

Light jet aircraft capable of carrying 6–10 passengers over short and medium ranges find great demand in the present and future civil aviation market. Light jets often cruise at high subsonic Mach numbers enabling greater range. Consequently, the drawback associated with high Mach number flight is that it leads to the deterioration of the aircraft’s aerodynamic characteristics. Aerodynamic optimization of a typical transonic wing was performed to reduce the wing cruise drag coefficient at a fixed Mach number and lift coefficient. A surrogate-based 2D optimization of wing sections was performed. Objective function values were obtained by solving Reynolds-averaged Navier–Stokes equations. The optimal baseline wing was constructed using the optimal airfoil sections obtained from the 2D optimization. Subsequently, parametric studies were conducted using vortex lattice method to arrive at an optimal wing by varying the wing twist distribution. The optimal airfoils showed 2% and 4% reduction in drag coefficient compared to the corresponding baseline airfoils. The optimal wing showed about 4% reduction in drag coefficient compared to the baseline wing.

Direct simulations are carried out to investigate the influence of unsteady heat flux transfer on transonic shock-boundary layer interaction; for flow past SHM-1 airfoil at a free-stream Mach number $$M_{\infty }$$ M ∞ = 0.72 and angle of attack $$\alpha = 0.38^{\circ }$$ α = 0 . 38 ∘ . Flux is added in a periodic manner through a region $$(8{-}18\% \; of \;the \;chord)$$ ( 8 - 18 % o f t h e c h o r d ) located on the suction side of the airfoil, with multiple values of exciter time period $$(T_{\text {e}}=2,4)$$ ( T e = 2 , 4 ) considered for our simulation. We show that addition of unsteady heat flux delayed shock formation, along with significant modifications in it’s structure. The time-averaged $$C_{\text {p}}$$ C p distributions revealed a shift in the shock towards the aft, by approximately 5% of the chord; along with an increased lift near the trailing edge, suggesting a nose-down stabilizing influence. Primarily, it is noted that the additional heat flux resulted in an overall increase of the aerodynamic efficiency (lift to drag ratio) by approximately $$10\%$$ 10 % .

This paper presents the development of a robust discrete adjoint approach for accurate computation of shape sensitivities in three-dimensional inviscid compressible flows. The adjoint Euler solver is generated by applying algorithmic differentiation techniques to the underlying primal solver. The novelty of the proposed framework is that the geometry subroutine that computes cell volumes, surface areas and normals is integrated to the subroutine that performs the primal fixed point scheme so that the adjoint code directly yields the desired shape sensitivities. The applicability of the developed adjoint approach is demonstrated on ONERA-M6 wing test case. The consistency and accuracy of the adjoint solver are assessed by comparing the adjoint shape sensitivities with the values from finite differences and tangent linear code. Numerical results show that the adjoint residual inherits the asymptotic rate of convergence of the primal residual.

This work presents the robust stability analysis of 2DoF airfoil by including various uncertainties. These uncertainties arise due to several factors such as modeling and manufacturing errors as well as disturbances in the flight conditions. The approach adopted to study the uncertain aeroelastic system is based on the structured singular value (µ-method). In this approach, the aeroelastic system is formulated in a robust stability framework by parameterizing around dynamic pressure and introducing uncertainties in the system parameters to account for errors and disturbances. This results in the perturbed aeroelastic system which is then represented using Linear Fractional Transformation (LFT). Then, the nominal and robust stability analysis of the perturbed aeroelastic system is carried out using µ method. In this work, first the validation of µ method is done for 2DoF airfoil with quasi-steady aerodynamics having uncertainties in the structural and aerodynamic properties. Further, the robust flutter boundary of 2DoF airfoil with Theodorsen’s unsteady aerodynamics is studied using µ method in the presence of stiffness, damping, and aerodynamic uncertainties.

Low Radar Cross Section (RCS) is one of the important requirements of aircraft design for stealth. Rotating engine face components are one of the major sources of radar reflection and hence need special attention. To this end, present day aircraft air intake ducts are designed to hide the rotating components from the radar signal by incorporating multiple bends. This type of intake ducts ensures that there is no direct line of sight from the entrance of the duct to engine face components and are generally called as Serpentine Ducts. Design of serpentine duct, therefore involves sharp bends that are likely to cause flow separation and consequent instability and distortion in the flow. In order to realize optimal performance from the engine, it is necessary to reduce the losses and minimize distortion at the Aerodynamic Interface Plane (AIP) that is near the outlet of the duct. This requires use of flow control techniques to alleviate the effects of flow separation. A Serpentine duct described in (Hamstra et al. in Active inlet flow control technology demonstration. ICAS, 2000 [1]) was selected for the present study, as the geometry details and experimental results were reported in the paper. Based on this, a duct was created and CFD flow simulations were performed. Results so obtained were compared with the test results in order to establish a baseline duct. Subsequently, passive vortex generators of trapezoidal shape were introduced and CFD simulations were carried out. The results, in comparison with the base design, indicate an enhancement in the flow uniformity at the AIP although with a 3% reduction in the pressure recovery. Further course of study is also indicated.

Hot flow analysis was undertaken in a cartridge starter using NUMECA software. Cartridge starter provides high temperature gases with kinetic energy which impinges on to the rotor blades of the turbine at the initial stage of starting a gas turbine engine which sets gas turbine to rotate. Other starting mechanisms which can be used are air starter, electric motor or a small gas turbine engine. The cartridge starter assembly consists of gas generator section with perforated central core, vertical tube/pipe which bifurcates into two pipe lines, called short arm and long arm, which provide high velocity hot gases to the high pressure turbine rotor blades. These two pipelines form convergent–divergent nozzles. Computational fluid dynamics (CFD) analyses were carried out for nine operating points for the inlet pressure ranging from 90 to 157 bar. Analyses considered the cartridge starter gas flow domain only and burning model is not considered. Hexahedral cells are generated with cut cell approach using fine Hexpress software. The problem is considered as three-dimensional, steady, compressible with turbulent flow for the flow analysis. The aim of the project is to determine the mass flow rate, its distribution in two arms and the total pressure loss. Spalart–Allmaras turbulence model with extended wall function was chosen for this analysis. First analysis showed that the flow was unsteady at the exit plane of nozzle. Therefore, analysis was carried out by considering unsteadiness in the flow and averaged quantities are derived. Flow is found to diffuse in the divergent nozzle after a normal shock. Mass flow rate has reduced due to the formation of shock. It is also found that the total pressure loss has reduced near the perforated central core of cartridge starter.

Screech combustion instabilities are high frequency (>1000 Hz) transverse periodic oscillations driven by combustion and which are then manifested as large amplitude oscillations in the afterburner duct pressure, accompanied by the characteristic high-pitched audible tones. These screech instabilities which are detrimental to the engine are conventionally suppressed by embedding Helmholtz resonator arrays in the afterburner liner. This method has been found inadequate when mixed mode combustion instability oscillations occur and also when the frequencies of oscillation were lower. The design of practical Helmholtz resonator arrays is classified and so is not available in the open domain. Hence, it was necessary to evolve a robust design solution to mitigate screech combustion instabilities in an afterburner. In an afterburner, V-gutters are used as flame stabilizers. The high Reynolds number flow past a V-gutter array is dominated by the presence of vortices characterized by the Kelvin–Helmholtz instability, which is a convective flow instability related to the shear layers separating from the V-gutter lips and the Benard–von Karman instability which is related to the asymmetric vortex shedding of the flow in the flame holder wake. The shedding of von Karman vortices at non-reacting and near the blowout conditions, and the transition from a Kelvin–Helmholtz instability to that of a Bernard–von Karman instability during near flame blowout create conditions for the frequency to get locked-on to the duct transverse acoustic mode frequency; screech is triggered. Hence, a smart flame stabilization method which has the intrinsic property of preventing the lock-on between the frequency of the vortex shedding from the V-gutter and the duct transverse acoustic frequency was developed. The test rig with optically accessible critical zones around the V-gutter flame stabilizer had the capability to operate the afterburner model under simulated inlet conditions of pressure and temperatures. A FastCam SA-4 Photron high-speed camera was used in this experimental investigation and high-speed shadowgraph flow visualization studies were carried out to develop a comprehensive method of introducing an aerodynamic splitter plate concept through a ventilated V-gutter; mitigation of screech combustion instability has been demonstrated.

Aerodynamic rolling moment on a core-alone launch vehicle due to the presence of a wire tunnel and the associated roll dynamics has been studied. Computational Fluid Dynamics (CFD) simulations across the roll angles from $$\phi $$ ϕ = 0 $$^{\circ }$$ ∘ to 180 $$^{\circ }$$ ∘ and across the Mach number range indicated the wire tunnel to be the major cause of the rolling moment. A passive means of roll moment reduction has been proposed by adding dummy wire tunnels symmetrically around the vehicle. It was found that adding one dummy wire tunnel diagonally opposite to the existing wire tunnel did not reduce the peak rolling moment as the leeward wire tunnel is ineffective. However, adding two dummy wire tunnels reduced the rolling moment substantially. Addition of the third dummy wire tunnel was also helpful in reducing the rolling moment further, though marginally. In addition to the CFD studies, the maximum roll rates and roll errors of the different configurations have been compared through roll dynamic simulations. A novel linear superposition methodology has been proposed and validated to obtain the rolling moment coefficient for multiple wire tunnel configurations.

Wind tunnel tests on high-speed air-intake configuration needs an accurate simulation of mass flow through the intake ducts. This leads to the requirement of miniature mass flow control device which provides linear variation in the throat area. Here, an attempt has been made to design and develop new miniature mass flow control device to characterize the air-intake model in the 1.2 m wind tunnel. This paper describes the design and development aspect of miniature mass flow control unit, drive electronics and its utility in wind tunnel testing of air-intake models. Emphasis is placed on experimental results obtained from an electrically actuated plug which controls the critical flow area at the downstream end of the intake model.

This paper discusses the effect of variable inlet guide vanes (VIGV) on the performance of military engine fan. The variable inlet guide vanes (VIGV) are necessary in order to safely start up multi-stage axial compressors. It not only improves surge margin at off-design condition but also gives nearly flat efficiency for wider speed range. The present fan design is a transonic three-stage axial flow compressor without inlet guide vanes. The aerodynamic performance of fan was limited by issues of lower stability range at part speeds, and blade flutter was observed in engine testing. So, it was proposed to redesign the fan with VIGV for addressing the above-mentioned design issues. Fixed-flapping type variable inlet guide vanes are designed to improve the performance of present fan. The design concept of VIGV itself will be a new design and development in our country. As a result of introducing VIGV, aerodynamic performance of redesigned fan is improved in terms of better surge margin at speeds lower than 95% due to reduced inlet incidence on rotor—stage1. The baseline fan was having stall flutter at part speed 80–85%; it is expected to get eliminated due to reduced incidence and improved stall margin.

In this work, a low aspect ratio MAV fixed wing at a relatively low Reynolds number wherein the flow undergoes transition is analysed. The effectiveness of the correlation based transition model $$\gamma $$ γ - $$Re_\theta $$ R e θ SST proposed by Menter and Langtry (Correlation based transition modeling for unstuctured parallelized computational fluid dynamics codes. AIAA J 47:2894–2906 [7]) is brought out by making vis-a-vis comparison with the pure turbulence model SST (Turbulence, heat and mass transfer vol 4. Begell House Inc., pp 625–626 [6]). The transition model is able to handle separated flow transition and gives more insight to flow than the turbulence model. Some of the results depicting the transitional flows are presented and the superiority of the transitional model over the pure turbulence model is demonstrated.

Panel methods are known to be simple yet effective during initial design stages. With the advent of advanced CFD methods and high speed computers, there is a feeling that these methods are no longer needed. On the contrary, experience with many practical problems shows that their utility is significant. There are also some problems where panel methods may be even superior to CFD approaches. One such problem is that of flow field analysis of rotors in the presence of stationary fuselage. In this paper we make use of an unsteady panel method which is a simple extension of a steady panel method to assess the effect of presence of fuselage on the rotor wake. Another significant advantage an unsteady panel method offers is that it is particularly easy to ‘fly’ the rotor in the presence of fuselage and assess the wake. Typical flight scenarios include hover, forward flight, ascent and descent flights. Some specific applications include assessment of wake flow for weapon separation from helicopters and estimates of noise due to unsteady rotor loading.

The aerodynamic behaviour of the GAW-1 airfoil at high Reynolds number is analysed numerically using the Spalart-Allmaras (SA) turbulence model.Initially an inter-code comparison is carried out at $$Re=6 \times 10^6$$ R e = 6 × 10 6 and the aerodynamics characteristics obtained using the in-house flow solution code 3D-PURLES and the open source CFD tools SU2 and OpenFOAM are validated with the available measurement data. Based on this validation exercise, 3D-PURLES is used to study the effect of increasing Reynolds number (1, 3, 6, 9, 12 millions) on the aerodynamic characteristics.

The rolling moment characteristics of a launch vehicle (LV) arising out of joining an air-breathing cruise vehicle (CV) and a booster is investigated through CFD to explain experimental observed behaviour. The launch vehicle has a stabilising fin at the rear and is placed on bearing to have free roll with respect to body. The basic and control rolling moment of LV (without stabilising fin, to mimic freely rolling fin), obtained experimentally, is compared with the corresponding experimental data for CV and found that the values differ, contrary to that observed in literature. CFD is used to investigate the reason for this difference. Investigation in two representative Mach numbers 0.8 (subsonic) and 2.0 (supersonic), has revealed that the interstage flare in LV, has significant effect in modifying the rolling moment contribution of the CV fin. Apart from that, the booster attachment arm, launch shoes are also having some impact in modifying the rolling moment. It is evident that for a canard controlled vehicle, even if the stabilising fin is freely rolling, the after-body geometry has significant effect in dictating the aerodynamic characteristics, especially rolling moment.

This paper presents a multidisciplinary, multi-objective optimization approach towards low Radar Cross section (RCS) aerodynamic design of aerospace structures. As a typical example, design and analysis of aircraft intake duct for stealth behavior by integrating computational fluid dynamics (CFD) and computational Electromagnetics (CEM) has been demonstrated. The design constraints (inlet area, throat area, exit area, and diameter) are calculated based on the RAE M2129 diffuser and subsonic flow condition with Mach number 0.8 is considered at the Indian standard atmospheric conditions (ISA_SL + 15). Inlet shaping parameters for intake are modified based on super ellipse equation by retaining the area as constant. Shaping parameter samples have been generated by limiting major axis to maximum length and minor axis to minimum length using MATLAB code. Based on each curvature parameter, three geometries were opted and CAD models are generated from sample space. For these geometries, CFD and CEM analysis has been performed and corresponding pressure recovery and RCS at 10 GHz (X-band) has been estimated. The CFD-CEM performance analysis has been presented for the optimized intake duct geometry. Particle swarm optimization (PSO) in-conjunction with CFD solver and CEM developed indigenous RCS solver has been used for optimization of the designed duct towards stealth characteristics.

The new aviation policy in place has given much impetus to local connectivity that has been re-emphasized with the launch of UDAN program. Certainly, this program will propel introduction of a significant number of propeller-driven aircraft suitable for short hauls. It is also well known that propeller-driven aircraft can be far more fuel efficient. In this context, CSIR-NAL initiated development of 14-seat Saras, a twin engine propeller-driven aircraft in pusher configuration. Being essentially an ab-initio design, one needs to understand complex aerodynamics that results from a pusher configuration. There are a number of technical issues such as propeller-fuselage interaction, propeller induced noise, power-on drag that need to be understood for design inputs. It is also known that full blown CFD methods for such analysis are still not very mature. Even if they are, they consume enormous computing power and clock-time to provide meaningful design inputs. In this paper, we present application of NAL’s unsteady Panel Code (Unsteady panel method analysis. PD-CTFD/2016/1009, CSIR National Aerospace Laboratories, Bangalore [1]) for the analysis of Hartzel propeller for a combination blade setting and advance ratios. This is an initial step towards more complex analysis wherein fuselage and other components can also be added in a much simpler fashion. Results indicate that there is a good confidence in this approach and as such one can generate significant design data at initial design stage.

Reynolds Averaged Navier–Stokes (RANS) simulations have been performed over a civil transport aircraft at Mach number 0.17. The present work aims to understand the effect of various Reynolds number on the aerodynamic performance of aircraft since the Reynolds number is considered as a most important parameter in fluid dynamics. The current study also helps to identify the various hot spots for higher drag so that the design team can focus on these specific areas to minimise the total drag. The present CFD solver uses a Roe scheme for convective flux discretisation and Spallart-Allmaras turbulence model for eddy viscosity computations. Calculations showed that with an increase in Reynolds number, the maximum lift coefficient increases and minimum drag coefficient decreases. The paper also highlights the sectional Cp plots and pressure contours along with the velocity streamlines at different span-wise stations of an aircraft wing. The strength of vortex over a midboard flap gradually decreases with the increase in Reynolds number. The drag of various aircraft components is also shown in the form of Pie chart so that the major drag contributing parts can be identified.

An efficient and economical in-house code WISe (Weapon Integration and Separation) has been developed to predict the trajectory of stores (1000 lb bomb and 250 kg bomb) released from an Indian fighter aircraft. First the stores are released in isolated mode and validated with a commercial code CFD++ and flight test data. After successful validation, the in-house code is extended to predict the trajectory of 1000 lb bombs when they are mounted on the mid-board and inboard stations on either sides of the Fighter Aircraft. The bombs are released in the following sequence—mid-board (left) followed by mid-board (right) followed by inboard (left) followed by inboard (right). Time interval considered between subsequent releases is 100 ms. A heavy-duty ejector release unit (ERU) has been used for inboard bombs whereas for mid-board bombs a light duty ejector has been used. The histories of store positions, orientation and miss distance are presented.

The occurrence of shock wave over the typical launch vehicle model with conical heat shield at a free-stream Mach number of 0.95 was studied using flow visualization techniques. Both oil flow pattern, as well as shadowgraph image techniques, have been used to trace the flow features over the model. The study was conducted at a various angle of attack from 0 $${^{\circ }}$$ ∘ to 4 $${^{\circ }}$$ ∘ in steps of 1 $${^{\circ }}$$ ∘ . Instantaneous shadowgraph images of shock wave pattern over the surface of the conical heat shield region were captured using a highspeed camera. The image shows a series of compression weak waves which forms a standing normal shock wave on the heat shield region at angles of attack of 0 $${^{\circ }}$$ ∘ , 1 $${^{\circ }}$$ ∘ , and 2 $${^{\circ }}$$ ∘ . Further increase in angle of attack, the normal shock wave splits into lambda-shaped shock pattern. At an angle of attack 3 $${^{\circ }}$$ ∘ and 4 $${^{\circ }}$$ ∘ , shock-induced separation and reattachment process result in shock oscillations along the surface of constant heat shield cylindrical region. The frequency of shock wave oscillations was found to increase with the angle of attack from 3 $${^{\circ }}$$ ∘ to 4 $${^{\circ }}$$ ∘ . The oil flow pattern clearly evident that the average distance covered during the shock oscillations.

The design of air intake systems for modern fifth generation fighter aircraft is of great significance as it directly affects the stealth characteristics of the aircraft. The major sources of Radar Cross Section (RCS) in an aircraft are its cavities such as air intake, diverter passage, weapon bays, etc. A diverterless air intake provides two fold advantages in supersonic military aircraft. In comparison with traditional intake with boundary layer diverter, a diverterless intake offers better aerodynamic performance at high speeds. It also gets rid of the diverter passage and the bump ahead of the entry to the intake provides additional shielding to the incident radar waves, thereby reduces RCS of the aircraft. It also eliminates the need to have moving surfaces used for efficient compression at high supersonic speeds. This paper presents the design approach of a Diverterless Supersonic Intake (DSI) on a generic stealth fighter aircraft and performance comparison with a conventional intake with boundary layer diverter.

This paper deals with validation of predicted trajectory of a BVR missile launched from an Indian fighter aircraft with flight test results. The trajectory of the missile is predicted using an in-house code WISe (Weapon Integration and Separation). The code works based on decay factorization scheme, where the assumption is that the missile aerodynamic interactions with the parent aircraft during separation process is dependent on missile’s horizontal displacement. In the absence of quantitative flight test trajectory data, the images of trajectories of the missile captured at various instants from the chase aircraft during test are compared with the predicted trajectories (pictures) of missile at corresponding time steps. The paper also contains generation of static plume using a commercial code Fluent. The missile plume path along with its trajectory is modeled from the static plume using two different approaches (Gleissl and Deslandes in simulation of missile plumes for aircraft store compatibility assessments, [1])—(i) Flashlight method and (ii) Particle ejection method. An assessment of plume ingestion into the air intake has been carried out and compared with flight test data obtained from temperature sensor. After successful validation, the rigid plume obtained from flashlight method and flexible plume from particle ejection method is also used to predict the static temperature at defined section of a Laser Guided Bomb (LGB) seeker.

This paper presents the details about lateral-directional and longitudinal control of a tailless aircraft for landing condition. Tailless aircraft configurations are specifically designed for stealth and the absence of a vertical fin aids in low visibility to RADAR. For such configurations, yaw control becomes a major challenge. Usage of control surfaces, known as spoilers, specifically to obtain yaw-controllability is one of the approaches. For the considered tailless aircraft configuration, it was observed that due to deflection of spoilers, the effectiveness of elevons was significantly hampered. This decrease in effectiveness demands higher deflection of elevons which is not preferred due to actuator requirements. Another approach to tackle the yaw-controllability would be to use the existing elevons present in aircraft. This approach is discussed and analyzed in this paper. Critical landing condition have been chosen as the design point to cater for, namely: (a) to generate sufficient yawing, rolling and pitching moments to restore the aircraft (b) by using minimum deflection angles to control (c) to attain required CL at landing condition. Few results of this study are presented here.