Uncertainty Quantification in Helicopter Performance Using Monte Carlo Simulations

The effect of structural and aerodynamic uncertainties on the performance predictions of a helicopter is investigated. An aerodynamic model based on blade element and momentum theory is used to predict the helicopter performance. The aeroelastic parameters, such as blade chord, rotor radius, two-dimensional lift-curve slope, blade profile drag coefficient, rotor angular velocity, blade pitch angle, and blade twist rate per radius of the rotor, are considered as random variables. The propagation of these uncertainties to the performance parameters, such as thrust coefficient and power coefficient, are studied using Monte Carlo Simulations. The simulations are performed with 100,000 samples of structural and aerodynamic uncertain variables with a coefficient of variation ranging from 1 to 5%. The scatter in power predictions in hover, axial climb, and forward flight for the untwisted and linearly twisted blades is studied. It is found that about 20-25% excess power can be required by the helicopter relative to the determination predictions due to uncertainties.

Effect of gas injection on drag and surface heat transfer rates for a 30 degrees semi-apex angle blunt body flying at Mach 5.75

Effect of coolant gas injection in the stagnation region on the surface heat transfer rates and aerodynamic drag for a large angle blunt body flying at hypersonic Mach number is reported for two stagnation enthalpies. A 60° apex-angle blunt cone model is employed for this purpose with air injection at the nose through a hole of 2mm diameter. The convective surface heating rates and aerodynamic drag are measured simultaneously using surface mounted platinum thin film sensors and internally mounted accelerometer balance system, respectively. About 35–40% reduction in surface heating rates is observed in the vicinity of stagnation region whereas 15–25% reduction in surface heating rates is felt beyond the stagnation region at stagnation enthalpy of 1.6MJ/kg. The aerodynamic drag expressed in terms of drag coefficient is found to increase by 20% due to the air injection.

A new 6-component accelerometer force balance for short duration ground testing facilities

A new six-component accelerometer force balance is developed and used in the HST2 shock tunnel of Indian Institute of Science. Aerodynamic forces and moments for a hypersonic slender body measured using this balance system at a free stream Mach number of 5.75 and Reynolds number of 1.5 million and stagnation enthalpy of 1.5 and 2 MJ/kg are presented. These measured values compare well with the theoretical values estimated using modified Newtonian theory.

A novel free floating accelerometer force balance system for shock tunnel applications

In order to overcome the interference of the model mounting system with the external aerodynamics of the body during shock tunnel testing, a new free floating internally mountable balance system that ensures unrestrained model motion during testing has been designed, fabricated and tested. Minimal friction ball bearings are used for ensuring the free floating condition of the model during tunnel testing. The drag force acting on a blunt leading edge flat plate at hypersonic Mach number has been measured using the new balance system. Finite element model (FEM) and CFD are exhaustively used in the design as well as for calibrating the new balance system. The experimentally measured drag force on the blunt leading edge flat plate at stagnation enthalpy of 0.7 and 1.2 MJ/kg and nominal Mach number of 5.75 matches well with FEM results. The concept can also be extended for measuring all the three fundamental aerodynamic forces in short duration test facilities like free piston driven shock tunnels.

Pressure-sensitive paint on a truncated cone in hypersonic flow at incidences

The flow over a truncated cone is a classical and fundamental problem for aerodynamic research due to its three-dimensional and complicated characteristics. The flow is made more complex when examining high angles of incidence. Recently these types of flows have drawn more attention for the purposes of drag reduction in supersonic/hypersonic flows. In the present study the flow over a truncated cone at various incidences was experimentally investigated in a Mach 5 flow with a unit Reynolds number of 13.5�10 6m -1. The cone semi-apex angle is 15° and the truncation ratio (truncated length/cone length) is 0.5. The incidence of the model varied from -12° to 12° with 3° intervals relative to the freestream direction. The external flow around the truncated cone was visualised by colour Schlieren photography, while the surface flow pattern was revealed using the oil flow method. The surface pressure distribution was measured using the anodized aluminium pressure-sensitive paint (AA-PSP) technique. Both top and sideviews of the pressure distribution on the model surface were acquired at various incidences. AA-PSP showed high pressure sensitivity and captured the complicated flow structures which correlated well with the colour Schlieren and oil flow visualisation results. © 2012 Elsevier Inc.

PIV study on the dimple mid-plane of a narrow rectangular channel with dimples applied to one wall

This paper presents an investigation of the fluid flow in the fully developed portion of a rectangular channel (Aspect Ratio of 2) with dimples applied to one wall at channel Reynolds numbers of 20,000, 30,000, and 40,000. The dimples are applied in a staggered-row, racetrack configuration. Results for three different dimple geometries are presented: a large dimple, small dimple, and double dimple. Heat transfer and aerodynamic results from preceding works are presented in Nusselt number and friction factor augmentation plots as determined experimentally. Using particle image velocimetry, the region near the dimple feature is studied in detail in the location of the entrainment and ejection of vortical packets into and out of the dimple; the downstream wake region behind each dimple is also studied to examine the effects of the local flow phenomenon that result in improved heat transfer in the areas of the channel wall not occupied by a feature. The focus of the paper is to examine the secondary flows in these dimpled channels in order to support the previously presented heat transfer trends. The flow visualization is also intended to improve the understanding of the flow disturbances in a dimpled channel; a better understanding of these effects would lead the development of more effective channel cooling designs. Copyright © 2011 by ASME.

A comparative study of dragonfly inspired flapping wings actuated by single crystal piezoceramic

A dragonfly inspired flapping wing is investigated in this paper. The flapping wing is actuated from the root by a PZT-5H and PZN-7%PT single crystal unimorph in the piezofan configuration. The nonlinear governing equations of motion of the smart flapping wing are obtained using the Hamilton's principle. These equations are then discretized using the Galerkin method and solved using the method of multiple scales. Dynamic characteristics of smart flapping wings having the same size as the actual wings of three different dragonfly species Aeshna Multicolor, Anax Parthenope Julius and Sympetrum Frequens are analyzed using numerical simulations. An unsteady aerodynamic model is used to obtain the aerodynamic forces. Finally, a comparative study of performances of three piezoelectrically actuated flapping wings is performed. The numerical results in this paper show that use of PZN-7%PT single crystal piezoceramic can lead to considerable amount of wing weight reduction and increase of lift and thrust force compared to PZT-5H material. It is also shown that dragonfly inspired smart flapping wings actuated by single crystal piezoceramic are a viable contender for insect scale flapping wing micro air vehicles.

Theoretical-experimental study of shock wave-assisted metal forming process using a diaphragmless shock tube

The use of high-velocity sheet-forming techniques where the strain rates are in excess of 10(2)/s can help us solve many problems that are difficult to overcome with traditional metal-forming techniques. In this investigation, thin metallic plates/foils were subjected to shock wave loading in the newly developed diaphragmless shock tube. The conventional shock tube used in the aerodynamic applications uses a metal diaphragm for generating shock waves. This method of operation has its own disadvantages including the problems associated with repeatable and reliable generation of shock waves. Moreover, in industrial scenario, changing metal diaphragms after every shot is not desirable. Hence, a diaphragmless shock tube is calibrated and used in this study. Shock Mach numbers up to 3 can be generated with a high degree of repeatability (+/- 4 per cent) for the pressure jumps across the primary shock wave. The shock Mach number scatter is within +/- 1.5 per cent. Copper, brass, and aluminium plates of diameter 60 mm and thickness varying from 0.1 to 1 mm are used. The plate peak over-pressures ranging from 1 to 10 bar are used. The midpoint deflection, circumferential, radial, and thickness strains are measured and using these, the Von Mises strain is also calculated. The experimental results are compared with the numerical values obtained using finite element analysis. The experimental results match well with the numerical values. The plastic hinge effect was also observed in the finite element simulations. Analysis of the failed specimens shows that aluminium plates had mode I failure, whereas copper plates had mode II failure.

Vaporization dynamics of functional droplets in a hot laminar air jet

The vaporization characteristics of pendant droplets of various chemical compositions (like conventional fuels, alternative fuels and nanosuspensions) subjected to convective heating in a laminar air jet have been analyzed. Different heating conditions were achieved by controlling the air temperature and velocity fields around the droplet. A hybrid timescale has been proposed which incorporates the effects of latent heat of vaporization, saturation vapor pressure and thermal diffusivity. This timescale in essence encapsulates the different parameters that influence the droplet vaporization rate. The analysis further permits the evaluation of the effect of various parameters such as surrounding temperature, Reynolds number, far-field vapor presence, impurity content and agglomeration dynamics (nanosuspensions) in the droplet. Flow visualization has been carried out to understand the role of internal recirculation on the vaporization rate. The visualization indicates the presence of a single vortex cell within the droplet on account of the rotation and oscillation of the droplet due to aerodynamic load. External heating induced agglomeration in nanofluids leads to morphological changes during the vaporization process. These morphological changes and alteration in vaporization behavior have been assessed using high speed imaging of the diameter regression and Scanning Electron Microscopy images of the resultant precipitate. (C) 2012 Elsevier Ltd. All rights reserved.

Breakup and coalescence characteristics of a hollow cone swirling spray

This paper deals with an experimental study of the breakup characteristics of water emanating from hollow cone hydraulic injector nozzles induced by pressure-swirling. The experiments were conducted using two nozzles with different orifice diameters 0.3 mm and 0.5 mm and injection pressures (0.3-4 MPa) which correspond to Rep = 7000-26 000. Two types of laser diagnostic techniques were utilized: shadowgraph and phase Doppler particle anemometry for a complete study of the atomization process. Measurements that were made in the spray in both axial and radial directions indicate that both velocity and average droplet diameter profiles are highly dependent on the nozzle characteristics, Weber number and Reynolds number. The spatial variation of diameter and velocity arises principally due to primary breakup of liquid films and subsequent secondary breakup of large droplets due to aerodynamic shear. Downstream of the nozzle, coalescence of droplets due to collision was also found to be significant. Different types of liquid film breakup were considered and found to match well with the theory. Secondary breakup due to shear was also studied theoretically and compared to the experimental data. Coalescence probability at different axial and radial locations was computed to explain the experimental results. The spray is subdivided into three zones: near the nozzle, a zone consisting of film and ligament regime, where primary breakup and some secondary breakup take place; a second zone where the secondary breakup process continues, but weakens, and the centrifugal dispersion becomes dominant; and a third zone away from the spray where coalescence is dominant. Each regime has been analyzed in detail, characterized by timescale and Weber number and validated using experimental data. (C) 2012 American Institute of Physics. http://dx.doi.org/10.1063/1.4773065]

Nonlinear dynamic analysis of dragonfly inspired piezoelectrically driven flapping and pitching wing

The nonlinear equations for coupled elastic flapping-twisting motion of a dragonfly in- spired smart flapping wing are used for a flapping wing actuated from the root by a PZT unimorph in the piezofan configuration. Excitation by the piezoelectric harmonic force generates only the flap bending motion, which in turn, induces the elastic twist motion due to interaction between flexural and torsional vibrations modes. An unsteady aerodynamic model is used to obtain the aerodynamic forces. Numerical simulations are performed using a wing whose size is the same as the dragonfly Sympetrum Frequens wing. It is found that the value of average lift reaches to its maximum when the smart flapping wing is excited at a frequency closer to the natural frequency in torsion. Moreover, consideration of the elastic twisting of flapping wing leads to an increase in the lift force. It is also found that the flapping wing generates sufficient lift to support its own weight and carry a small pay- load. Therefore, the piezoelectrically actuated smart flapping wing based on the geometry of Sympetrum Frequens wing and undergoing flapping-twisting motions may be considered as a potential candidate for use in MAV applications.

Tuning of the extended concentric tube resonators

It has been shown recently that the acoustic performance of the extended tube expansion chambers can be improved substantially by making the extended inlet and outlet equal to half and quarter chamber lengths, duly incorporating the end corrections due to the evanescent higher order modes that would be generated at the discontinuities. Such chambers however suffer from the disadvantages of high back pressure and generation of aerodynamic noise at the area discontinuities. These two disadvantages can be overcome by means of a perforated bridge between the extended inlet and extended outlet. This paper deals with design or tuning of these extended concentric tube resonators.

Topology optimization-based design of a compliant aircraft compliant aircraft wing for morphing leading and trailing edges

In this paper, we present a methodology for designing a compliant aircraft wing, which can morph from a given airfoil shape to another given shape under the actuation of internal forces and can offer sufficient stiffness in both configurations under the respective aerodynamic loads. The least square error in displacements, Fourier descriptors, geometric moments, and moment invariants are studied to compare candidate shapes and to pose the optimization problem. Their relative merits and demerits are discussed in this paper. The `frame finite element ground structure' approach is used for topology optimization and the resulting solutions are converted to continuum solutions. The introduction of a notch-like feature is the key to the success of the design. It not only gives a good match for the target morphed shape for the leading and trailing edges but also minimizes the extension of the flexible skin that is to be put on the airfoil frame. Even though linear small-displacement elastic analysis is used in optimization, the obtained designs are analysed for large displacement behavior. The methodology developed here is not restricted to aircraft wings; it can be used to solve any shape-morphing requirement in flexible structures and compliant mechanisms.

Transition and acoustic response of recirculation structures in an unconfined co-axial isothermal swirling flow

This paper reports the first observations of transition from a pre-vortex breakdown (Pre-VB) flowreversal to a fully developed central toroidal recirculation zone in a non-reacting, double-concentric swirling jet configuration and its response to longitudinal acoustic excitation. This transition proceeds with the formation of two intermediate, critical flow regimes. First, a partially penetrated vortex breakdown bubble (VBB) is formed that indicates the first occurrence of an enclosed structure as the centre jet penetration is suppressed by the growing outer roll-up eddy; resulting in an opposed flow stagnation region. Second, a metastable transition structure is formed that marks the collapse of inner mixing vortices. In this study, the time-averaged topological changes in the coherent recirculation structures are discussed based on the non-dimensional modified Rossby number (Ro(m)) which appears to describe the spreading of the zone of swirl influence in different flow regimes. Further, the time-mean global acoustic response of pre-VB and VBB is measured as a function of pulsing frequency using the relative aerodynamic blockage factor (i.e., maximum radial width of the inner recirculation zone). It is observed that all flow modes except VBB are structurally unstable as they exhibit severe transverse radial shrinkage (similar to 20%) at the burner Helmholtz resonant modes (100-110 Hz). In contrast, all flow regimes show positional instability as seen by the large-scale, asymmetric spatial shifting of the vortex core centres. Finally, the mixing transfer function M (f) and magnitude squared coherence lambda(2)(f) analysis is presented to determine the natural couplingmodes of the system dynamic parameters (u', p'), i.e., local acoustic response. It is seen that the pre-VB flow mode exhibits a narrow-band, low pass filter behavior with a linear response window of 100-105 Hz. However, in the VBB structure, presence of critical regions such as the opposed flow stagnation region alters the linearity range with the structure showing a response even at higher pulsing frequencies (100-300 Hz). (C) 2013 AIP Publishing LLC.