Subsonic Range Research Articles

Experimental and Numerical Analysis Over Power Series Nose Cone

While designing aerospace vehicles, one of the most important factors to consider is drag. Drag force opposes the vehicle and causes it to slow down, which in turn requires more fuel to maintain flight. The vehicle nose cone plays a vital role in minimizing the drag during flight. The shape of the vehicle's body, as well as the fluid characteristics and orientation, can all have a significant impact on drag. To determine the most effective nose cone design, a study was conducted to test the aerodynamic performance of power series nose cone profiles using both experimental and computational techniques. All nose cone profiles with the same L/D ratio were assessed at 25 m/s speed with no angle of attack. The experimental data, including pressure, velocity, and drag, were then compared with computational data to ensure the accuracy of the wind tunnel experiments. The primary factor in determining the best shape for the subsonic flow range among all the nose cone profiles was the drag coefficient. The nose cone with a power of 0.25 was found to have the lowest drag coefficient at low subsonic speeds, making it the best-performing nose cone profile among those investigated. During take-off, drag is the most undesirable factor, as it can significantly impact the fuel efficiency of the vehicle. Therefore, it is crucial to have the least amount of drag possible. In this case, the nose cone with a power of 0.25 can be used in the initial stages of the rocket to minimize drag and maximize fuel efficiency.

Design and Analysis of Wing Tip Twist on a Forward Swept Wing Using Prandtl-d Wing Concept

Our project is all about the implementation of a wing tip twist for Prandtl-D wing to a Forward swept wing. The innovative idea is that the Prandtl-D wing is used only for a glider till date whereas, we implied this concept for a commercial aircraft. The phenomena of twisting the tip of the wing leads to the change in geometry which results in drag reduction and increase in lift to drag ratio. With the help of a reference project, a wing geometry is chosen and the model is scaled down according to our test section. By using CATIA V5 R20, our model is drawn and analyzed with the help of ANSYS Fluent 16.0. The geometry with the best result is 3-D printed and tested in the low speed subsonic wind tunnel. Both the analytical and experimental results are validated. In general, the forward swept wing aircraft possess good aerodynamic performance in subsonic range. We look into that whether our project increases the efficiency or not.

Finite energy traveling waves for the Gross-Pitaevskii equation in the subsonic regime

In this paper we study the existence of finite energy traveling waves for the Gross-Pitaevskii equation. This problem hasdeserved a lot of attention in the literature, but the existence of solutions in the whole subsonic range was a standing open problem till the work of Mari\\c\{s\} in 2013. However, such result is valid only in dimension 3 and higher. In this paperwe first prove the existence of finite energy traveling waves for almost every value of the speed in the subsonic range. Our argument works identically well in dimensions 2 and 3.

DYNAMIC MODEL OF THE DURATION OF GASEOUS ENVIRONMENT PUMPING FROM A LIMITED VOLUME

An analytical equation for the duration of air pumping from a limited volume has been developed. The equation of the mechanical energy of air movement takes into account the law of mass conservation for a gas in a controlled volume, the work of the energy of air movement and the work to overcome air friction. Gaseous medium is Newtonian. The duration of pumping (filling) the limited volume of the pneumatic chamber with air was calculated by comparing the mass flow per second and the increase in mass, as a differential of the change in air density. The mathematical model enables the simulation of air pumping time depending on pressure, as a density parameter and at different Mach numbers in the subsonic range. The K1* proportionality coefficient, which characterizes the ratio of the dynamic force of gas mass displacement to the static pressure relative to the diameter of the air pipeline, is proposed as a criterion for evaluating the dynamics of the flow. It should be noted that the analytical dependence works for Newtonian media and Mach numbers of M < 1, the gas flow is caused by the pressure difference, the gas itself is limited by a chamber space characterized by a volume as design parameter of structure.

Investigation of suddenly expanded flows at subsonic Mach numbers using an artificial neural networks approach.

The purpose of this study is to explore two concepts: first, the use of artificial neural networks (ANN) to forecast the base pressure (β) and wall pressure (ω) originating from a suddenly expanded flow field at subsonic Mach numbers. Second, the implementation of Garson approach to determine the critical operating parameters affecting the suddenly expanded subsonic flow process in the subsonic range. In a MATLAB environment, a network model was constructed based on a multilayer perceptron with an input, hidden, and output layer. The network input parameters were the Mach number (M), nozzle pressure ratio (η), area ratio (α), length to diameter ratio (γ), micro jet control (ϵ), and duct location to length ratio (δ). The network output included two variables; base pressure (β) and wall pressure (ω). The ANN was trained and tested using the experimental data. The experimental results found that micro-jet controls were successful in increasing the base pressure for low Mach numbers and high nozzle pressure ratios. It was also found that the wall pressure was same for with and without micro jet control. The ANN predicted values agreed well with the experimental values, with average relative errors of less than 5.02% for base pressure and 6.71% for wall pressure. Additionally, with a relative significance of 32% and 43%, the nozzle pressure ratio and duct location to length ratio had the highest influence on the base pressure and wall pressure, respectively. The results demonstrate that the ANN model is capable of accurately predicting the pressure results, enabling theoretical foundation for research into pressure distribution in aerodynamic systems.

Investigation on the Reynolds Number Effect of a Flying Wing Model with Large Sweep Angle and Small Aspect Ratio

The effect of the Reynolds number on the flying wing common research model with a large sweep angle and small aspect ratio is studied by numerical simulation. The helicity cross-flow transition correction of the Langtry–Menter model is established in view of the deficient prediction accuracies of existing transition prediction models in simulating cross-flow transition, and the helicity parameters are calibrated to improve the prediction accuracy of the numerical method. The transition prediction method is verified by the test results of the standard model transition of the DLR-F5 wing and small-aspect-ratio flying wing. At the same time, the numerical method is effectively verified by comparing it to the aerodynamic/torque results of the standard model of the small-aspect-ratio flying wing. On this basis, the variation laws with the Reynolds number are analyzed for the aerodynamic and flow field characteristics of the flying wing common research model with a large sweep angle and small aspect ratio. The numerical simulation is mainly carried out in the high subsonic speed range (M = 0.9), and the simulated Reynolds number range is from 2 × 106 to 80 × 106, which includes the test Reynolds number and flight Reynolds number. The influence of turbulence on the surface flow pattern and drag characteristics of the small-aspect-ratio standard model is analyzed. Calculation results show that when the turbulence is greater than 0.4%, it has a considerable influence on the surface laminar flow range and total drag coefficient. The results reveal that the Reynolds number mainly affects the magnitude of friction in the cruise state (small angle of attack). The problem of the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio is further studied, and the Reynolds number is found to be greater than 10 × 106. The drag and lift coefficients after the approach are close to the Reynolds number in the self-aligning region of the flying wing standard model with a small aspect ratio.

Study on the Characteristics of Boundary Layer Flow under the Influence of Surface Microstructure

The energy consumption of a vehicle is closely related to the resistance it receives, and it is of great significance to study the drag reduction of a vehicle to promote energy conservation and emissions reductions. Boundary layer control drag reduction is mainly achieved by controlling the coherent structure in turbulence and reducing its burst intensity and frequency. It can be divided into an active control drag reduction and passive control drag reduction. In passive drag reduction, the advantages of the surface groove drag reduction are relatively obvious. In this paper, the large eddy simulation method is used to study the boundary layer flow with triangular groove and rectangular groove plates along the flow direction under subsonic flow, and to explore the influence of a surface micro-groove structure on the boundary layer flow. The simulation results show that the fluid inside the groove can be blocked by the triangular groove which can keep the low-velocity fluid at the bottom of the groove, and that it can increase the thickness of the viscous bottom layer as well as reduce the velocity gradient at the wall. The spanwise stress component of the Reynolds stress in the triangular groove boundary layer and the burst of turbulence on the wall are inhibited, and the spanwise flow in the boundary layer is blocked. In the subsonic range, about 10% shear force can be reduced because there are secondary vortices induced by the upper flow vortices at the top of the groove wall, and these secondary vortices can restrain the rising of the low-speed strip in the groove and reduce the burst of turbulence. The rectangular groove creates a weak blocking effect on the fluid inside the groove, which can only inhibit spanwise pulsation under subsonic speed. The wall shear stress cannot be reduced when the flow velocity is low, and it even increases.

Tonal Noise Suppression of an Underexpanded Orifice Jet Upon Impingement Over Corrugated Geometries

Abstract The current study investigates the noise generation of an orifice jet upon impingement over different corrugated geometries for different nozzle pressure ratios (NPRs). Semi-circular corrugations of different pitch lengths are considered for this study. A comparison of corrugated geometries is made with the flat plate. A standoff distance of 3.5 times jet diameter is considered during the experiment. The noise propagation at the far-field is observed using a far-field microphone. Several tones and their harmonics of the feedback loops established between the orifice exit and the plates are analyzed. The acoustic spectra show that the corrugated geometries emit less noise at subsonic speed compared to supersonic speed. A reduction of overall sound pressure level (OASPL) up to a range of 3–8 dB is observed for the corrugated plate compared to the base plate at the subsonic range. The tonal noise is reduced for the corrugated geometries at all NPRs. The wavelength of the primary tones is compared with the pitch length of semi-circular corrugations. Schlieren images show the presence of the acoustic feedback loop and standing waves near the impingement region for the flat plate. The flow field in between the semi-circular corrugations is analyzed by solving the large eddy simulation. The directivity study shows a reduction in OASPL value at the upstream direction at NPR 4 and 4.8 for the corrugated geometries.

Numerical investigation of flow over a hemispherical missile nose cone configuration in subsonic speed

Numerical investigation of flow around an hemispherical nose cone missile was simulated for very low speed flow conditions. The geometry creation, meshing of geometry, solving and post processing was done in ANSYS - Fluent. This research uses CFD at the subsonic range to analyze missiles with hemispherical nose cones at various angles of attacks from 0° to 16° with step of 4° deg and different velocity ranging from 5 m/s to 20 m/s. This study estimates the aerodynamics characteristics such as the coefficient of drag and lift. Besides, distribution of the coefficient of pressure over the surface of the hemispherical missile nose cones and their graphical result are ploted and presented.

Modelling the penetration of subsonic rigid projectile probes into granular materials using the cavity expansion theory

Typical subsurface investigation tools used on Earth are not applicable for subsurface exploration of the Moon and other extraterrestrial bodies due to payload limitations in space missions. Instead, light and compact subsonic projectile probes can be considered as an alternative subsurface investigation tool in such exploratory missions to overcome a host of challenges. Such probes can be launched from a lunar orbiter or lander to the surface of the Moon to provide the initial effective penetration from the impact. Here, we develop a model based on the spherical cavity expansion theory to predict the deceleration rate and final penetration depth of a rigid projectile probe into geological targets under the perpendicular subsonic impact. Two stress fields are assumed to propagate in the medium, plastic (near field) and elastic (far field), upon the impact. The stresses at the cavity wall are obtained by combining the Mohr–Coulomb failure criterion for the target failure (plastic region) considering two different assumptions for plastic wave propagation. Two field experiments are used to compare and assess the robustness of the proposed solutions in the subsonic range. Base on the simulation results and the experiments, it is concluded that the cavity expansion model considering the locked hydrostat assumption, with the modification here introduced for the volumetric strain, can provide us with a reasonable prediction of the projectile penetration, final penetration depth and the stresses on the probe. Thus, our proposed solution can be used as a benchmark for sophisticated and computationally-expensive numerical calculations.

Artificial intelligence-based stethoscope for the diagnosis of aortic stenosis

Abstract Background Although the stethoscope is used for auscultation in the diagnosis of valvular heart disease for more than 200 years, its diagnostic accuracy is limited and highly dependent on clinical expertise and acoustic range of the human ear. Purpose To develop an electronic stethoscope, based on artificial intelligence (AI), for the diagnosis of aortic stenosis (AS). Methods We developed an electronic stethoscope (VoqxTM, Sanolla) with subsonic capabilities and acoustic range of 0–2,000 Hz. Using the VoqxTM, we recorded heart sounds from 100 patients referred for echocardiography (derivation group), 50 with moderate or severe AS (aortic valve area (AVA) ≤1.5 cm2) and 50 without valvular disease, using the 5 standard auscultation points. An AI based supervised learning model was applied to the auscultation data from the first 100 patients, to construct a diagnostic algorithm that was then tested on a validation group (50 other patients, 25 with AS and 25 without AS). Results The derivation group included 50 patients with AS (age 78±9 years, 28 males, AVA=0.87±0.19 cm2, mean gradient 44±16 mmHg, 39 with severe AS) and 50 without AS (age 58±16 years, 39 males). The AI based algorithm using 1–4 auscultation position (all except the mitral position), correctly identified 47/50 AS patients (sensitivity 94%) and 49/50 patients without AS (specificity 98%), total accuracy 96% (Table). The algorithm was then applied to the validation group (AS: n=25, age 76±9 years,12 males, AVA=0.88±0.21 cm2, mean gradient 41±15 mmHg, 18 with severe AS; No AS: n=25, age = 59±18 years, 17 males). The algorithm correctly identified 21/25 AS patients (sensitivity 84%) and 24/25 patients without AS (specificity 96%), total accuracy 90% (Table 1). Conclusion Our initial findings show that an AI based stethoscope can accurately diagnose AS. AI based electronic auscultation is a promising new tool for automatic diagnosis of valvular heart disease. Funding Acknowledgement Type of funding sources: Private company. Main funding source(s): Sanolla Company Table 1

Deep penetration of low velocity metal penetrators into ice

The paper summarizes the results of research of the destruction of an ice block with cylindrical and spherical penetrators at low velocity (≤325 m/s). The behaviour of ice at high strain rates is described by an elastic–plastic model of continuum mechanics. Numerical modelling of penetration is performed with IMPACT computer code. Algorithms of splitting nodes and destroying elements in a Lagrangian numerical method were modified to solve problems of penetration and perforation. The fracturing is described by a deterministic fracture model. Crack paths are examined; the damage is predicted and compared with existing experimental results. It was found that in the subsonic range of initial velocities the penetration time did not exceed 0.3 ms. Examination of the penetrators’ shapes showed that they were not plastic deformed.

Liftoff height of turbulent diffusion flame in water-mist-laden coflowing air stream

In the present study, the enhancement of the liftoff and blowout of a propane/air coflowing diffusion flame by water mist was experimentally investigated in the subsonic range. When the fuel injection velocity Uf is low, an attached laminar diffusion flame is stabilized at the burner rim with a premixed flame base. An increase of Uf leads to the liftoff of the attached flame, and eventually to the blowout of the lifted flame. The upstream laminar flame disappears at the liftoff and a turbulent lifted flame is stabilized in the downstream region. The stability of the lifted flame deteriorates with the increase of the coflowing air velocity Ua and the water mist flow rate Qm. The stable lifted flame can be achieved by the dynamic balance of the local mean gas velocity Ug¯ and the mean turbulent flame propagation speed Up¯ of the premixed flame formed at base of the lifted flame, and Up¯ was estimated by the measurements of Ug¯ by a laser Doppler velocimeter. Up¯ decreases with Qm and is of the order of 1 m/s at the blowout independently of Ua and the mass fraction of water mist Y0. This value is nearly equal to 2.5SL0, where SL0 is the adiabatic laminar flame speed of a stoichiometric propane-air mixture. The liftoff height L increases both with Uf and Qm. For the case without water mist, a unique correlation has been proposed in the previous studies between L and Uf in a non-dimensional form using SL0. By extending this correlation to the diffusion flame in the water-mist-laden air stream, the laminar flame speed influenced by water mist was appraised and compared with the simulation previously reported. Resulting flame speed reduction is 65% of that predicted by numerical simulation, presumably due to the partial evaporation of water mist.

High subsonic flow of an organic vapor past a circular cylinder

A circular cylinder was tested in the cross-flow of an organic vapor (Novec™ 649) and of air over the subsonic (M < 0.4) and high subsonic (0.4 < M < 0.8) speed range in a continuously running pressurized closed-loop wind tunnel test facility. Time-averaged pressure measurements gave information on surface pressure distributions, and the corresponding drag and base pressure drag coefficients were obtained. Due to the charging of the wind tunnel, different values of the compressibility factor (0.876 < Z < 0.999) could be achieved for the organic vapor flow. This enabled in combination with the results for air an assessment of the impact of non-ideal gas dynamics on the form drag of a cylinder in the considered highly subsonic flow regime. The new experimental data were compared with available literature results. Changes in surface pressure distribution at higher subsonic velocities were identified and discussed. It was found that non-ideal gas effects did not strongly affect the overall drag. The variation of drag coefficient over the Mach number range was comparable with literature data for ideal-gas compressible flow, including shock-less and intermittent shock wave, and permanent shock wave flows regimes. At Mach 0.4, the flow of Novec™ 649 was in the shock-less regime and exhibited a pronounced dependency on the Reynolds number. An increase in drag was observed at Mach 0.6 which was attributed to the commencement of vortex shedding. Non-ideal thermodynamics only affected the flow locally and a reduction of the critical pressure coefficient in the high subsonic flow regime was observed in the surface pressure distribution. However, this mechanism did not alter significantly the overall drag behavior.Graphic abstractDrag coefficient CD against Re for several Mach numbers M and comparison with available literature results obtained for air (colored symbols indicate different Mach number clusters)

Improved Estimation of Air-to-Air Missile Aerodynamic Characteristics with Using CFD Method

The paper describes a solving procedure of the special Finite Volume Method (FVM) problem dealing with calculation of a flow field around air-to-air missile at subsonic, transonic and supersonic velocity range in order to obtain a coefficient of drag (cD). Some emphasis is put on numerical solution and fundamentals of FVM and use of a k-&amp;omega; turbulence model for simulation of subsonic and supersonic flow fieldsaround air-to-air missile AIM-9 Sidewinder version M. Obtained results of missile aerodynamics characteristics are verified using semi-empirical software and by analysis of flow fields.

Classical flutter analysis of composite wind turbine blades including compressibility

AbstractFor wind turbine blades with the increased slenderness ratio, flutter instability may occur at lower wind and rotational speeds. For long blades, at the flutter condition, relative velocities at blade sections away from the hub center are usually in the subsonic compressible range. In this study, for the first time for composite wind turbine blades, a frequency domain classical flutter analysis methodology has been presented including the compressibility effect only for the outboard blade sections, which are in the compressible flow regime exceeding Mach 0.3. Flutter analyses have been performed for the baseline blade designed for the 5‐MW wind turbine of NREL. Beam‐blade model has been generated by making analogy with the structural model of the prewisted rotating thin‐walled beam (TWB) and variational asymptotic beam section (VABS) method has been utilized for the calculation of the sectional properties of the blade. To investigate the compressibility effect on the flutter characteristics of the blade, frequency and time domain aeroelastic analyses have been conducted by utilizing unsteady aerodynamics via incompressible and compressible indicial functions. This study shows that with use of compressible indicial functions, the effect of compressibility can be taken into account effectively in the frequency domain aeroelastic stability analysis of long blades whose outboard sections are inevitably in the compressible flow regime at the onset of flutter.

Parametric transition from deflagration to detonation revisited: Spherical geometry

The paper is concerned with identification of the key mechanisms controlling deflagration-to-detonation transition in unconfined gaseous systems. The issue of thermal runaway triggered by positive feedback between the advancing flame and the flame-driven precompression wave is discussed in the context of expanding spherical flames. Depending on parameters of the system the transition may occur either within a subsonic or supersonic range of flame propagation. The paper is an extension of the previous study concerned with planar flames.

Error Estimation of Direction Finding and Determining Coordinates of Signal Source Using Combined Hydroacoustic Receivers

The paper deals with the possibility of direction finding and determining the coordinates of a source of acoustic signal emitted by a surface or underwater vehicle in the subsonic frequency range. For this purpose combined hydroacoustic receivers are used. During the first experiment, the estimates of the systematic component of direction finding error were obtained when the source emitted a tone or polyharmonic signal, and the direction finding results were averaged over a set of frequencies. The obtained results were used in the second experiment for estimating the error in determining the coordinates of the polyharmonic signal emitter consisting of three combined receivers (CR) which form the navigation base distributed in the horizontal plane.

Investigation of flow-field around a single generic planar fin using CFD

The purpose of this paper is to initiate a 3-D non-spinning semi-circular missile model with a single generic planar fin and perform a computational analysis on it to understand the flow pattern around the fin. A freestream computational fluid dynamics analysis was done in the subsonic and supersonic Mach range, in which, the fluid behaviour was investigated using a two-equation turbulence model. A structured mesh was adopted to visualize the flow pattern around the fin. The aerodynamic coefficients were calculated for the model, and the predicted values were compared with the previous experimental results as well as the numerical results. This paper attempts to present all the possible flow visualizations which might help in better understanding of flow around missiles having planar fins. This work also attempts to establish a turbulent computational model for a single planar fin missile model for subsonic to supersonic range.

Reduced order nonlinear aeroelasticity of swept composite wings using compressible indicial unsteady aerodynamics

Nonlinear dynamic aeroelasticity of composite wings in compressible flows is investigated. To provide a reasonable model for the problem, the composite wing is modeled as a thin walled beam (TWB) with circumferentially asymmetric stiffness layup configuration. The structural model considers nonlinear strain displacement relations and a number of non-classical effects, such as transverse shear and warping inhibition. Geometrically nonlinear terms of up to third order are retained in the formulation. Unsteady aerodynamic loads are calculated according to a compressible model, described by indicial function approximations in the time domain. The aeroelastic system of equations is augmented by the differential equations governing the aerodynamics lag states to derive the final explicit form of the coupled fluid-structure equations of motion. The final nonlinear governing aeroelastic system of equations is solved using the eigenvectors of the linear structural equations of motion to approximate the spatial variation of the corresponding degrees of freedom in the Ritz solution method. Direct time integrations of the nonlinear equations of motion representing the full aeroelastic system are conducted using the well-known Runge–Kutta method. A comprehensive insight is provided over the effect of parameters such as the lamination fiber angle and the sweep angle on the stability margins and the limit cycle oscillation behavior of the system. Integration of the interpolation method employed for the evaluation of compressible indicial functions at any Mach number in the subsonic compressible range to the derivation process of the third order nonlinear aeroelastic system of equations based on TWB theory is done for the first time. Results show that flutter speeds obtained by the incompressible unsteady aerodynamics are not conservative and as the backward sweep angle of the wing is increased, post-flutter aeroelastic response of the wing becomes more well-behaved.