Book of Abstracts - iitk.ac.in · Effect of air to liquid momentum flux ratio on primary jet...

First Aerospace Engineering Doctoral Students’ Symposium

May 12, 2014,

Department of Aerospace Engineering,

IIT Kanpur, India 208016.

Book of Abstracts

May 2014

Kanpur, India

The Symposium is organized by the doctoral students of the Department of Aerospace

Engineering, IIT Kanpur.

Advisory Committee:

1. Dr. Ashoke De

Convener, Department Post Graduate Committee

2. Dr. Sanjay Mittal

Head of the Department

3. Dr. C. Venkatesan

4. Dr. Debopam Das

Table of contents

Sl. no.

Abstract number Title Page no.

AERODYNAMICS

1 AEDSS-2014-101 Velocity and Density Measurements in a Circular Buoyant Helium Plume 1

2 AEDSS-2014-102 Velocity and Density field of compressible vortex rings 2

3 AEDSS-2014-103 Experimental Aerodynamics of a Butterfly mimicking Flapping Wings at Low Reynolds Number

3

4 AEDSS-2014-104 Viscous flow in a twin intake at supersonic speeds 4

5 AEDSS-2014-105 Experimental investigation of asymmetric pitching oscillations of a symmetric airfoil

5

6 AEDSS-2014-106 Velocity Dynamics of Suddenly Blocked Oscillatory Channel Flow 7

7 AEDSS-2014-107 Numerical Investigation of Clap and Fling Motion Using Immersed Boundary Lattice Boltzmann Method

8

8 AEDSS-2014-108 Normal interaction of a compressible vortex ring on the wall 10

9 AEDSS-2014-109 Free vibrations of a cylinder beyond the laminar regime 12

10 AEDSS-2014-110 Analysis of Aerodynamic Forces and Inflight Measurements Of An Ornithopter 14

AEROSPACE PROPULSION

11 AEDSS-2014-201 Emission Measurements from Bench Scale Aircraft Combustor Rig 16

12 AEDSS-2014-202 Large Eddy Simulation of Swirling Non Reactive Flow 18

13 AEDSS-2014-203 Effect of transverse periodic loading on an airfoil in a cascade 20

14 AEDSS-2014-204 Computation of Supersonic Flow Past Backward Facing Step in OpenFOAM 22

15 AEDSS-2014-205 Numerical Investigation of Soot Formation in Turbulent Diffusion Flame 24

16 AEDSS-2014-206 Effect of air to liquid momentum flux ratio on primary jet breakup in cross flow of air at atmospheric pressure

26

17 AEDSS-2014-206 Active Control of Hooting in Gas Turbine Engines 28

AEROSPACE STRUCTURES

18 AEDSS-2014-301 Addition of a lead lag damper model and change of root boundry conditions in comprehensive aeroelastic code and its effects on structural dynamics and blade loads

31

19 AEDSS-2014-302 Development and Structural Dynamic Analysis of Bio-inspired MAV Flapping Wings

33

20 AEDSS-2014-303 A-Posteriori Error Estimation for Non-Linear Problems 34

FLIGHT MECHANICS AND CONTROL

21 AEDSS-2014-401 Angle of Attack, Pitch Angle and Glide Angle Modeling at Various Thrust Inputs for a Powered Parachute Aerial Vehicle

35

22 AEDSS-2014-402 Flight Dynamic Modeling for Mini Helicopter with Stabilizer bar for Trim and Stability Analysis

39

23 AEDSS-2014-403 Influence of Main Rotor downwash on the Horizontal Tail and its effect on Trim of the Helicopter

41

24 AEDSS-2014-404 Estimation of Nonlinear Parameters from Simulated Data of an Aircraft

43

25 AEDSS-2014-405 Design, Instrumentation and Data Acquisition of 5 Degree of Freedom (5-DOF) Dynamic Test Rig

46

Aerospace Engineering Doctoral Students’ Symposium (AEDSS) IIT Kanpur, May 12-13, 2014

1

AEDSS-2014-101

Velocity and Density Measurements in a Circular Buoyant Helium Plume

Kuchimanchi K Bharadwaj

1,a, Debopam Das

1 and Pavan K Sharma

2

1Department of Aerospace Engineering, Indian Institute of Technology Kanpur, Kanpur

2Reactor Safety Division, Bhabha Atomic Research Center, Mumbai

(aCorresponding author, e-mail: [email protected])

Medium scale buoyancy dominated helium plumes are experimentally investigated in the present study. It

has been well known that buoyant helium plumes released into air can be used to simulate flames/pool

fires because of the similarity in the flow features and unsteady dynamics exhibited by them in the source

near field1,2

. In a certain parameter space (i.e. Reynolds number Re, Froude number Fr and Density ratio

S), these buoyant helium plumes exhibit pulsating instability in the near field, which is similar to the

puffing phenomenon observed in buoyant diffusion flames and pool fires. This instability manifests itself

as the periodic shedding of vortical structures from the plume source. This periodic shedding/puffing

determines the near field entrainment characteristics of these plumes2. The present study is carried out

with particular emphasis on the instability manifestation and ambient fluid entrainment in circular helium

plumes. The study of entrainment in such flows requires both velocity and helium concentration

measurements. Two-dimensional velocity field has been measured using Particle Image Velocimetry

(PIV). Helium concentration/density field has been measured using Planar Laser Induced Fluorescence

(PLIF), which is developed in-house. Dependence of puffing instability and ambient fluid entrainment on

various parameters will be discussed in the present study.

Figure 1: Velocity, vorticity and density fields at an instant during puffing cycle for helium plume at

Re=117, Fr=0.3 & S=0.14.

References:

1. Cetegen, B.M., Ahmed, T.A.: Experiments on the periodic instability of buoyant plumes and pool

fires. Combust. Flame 93, 157–184 (1993)

2. Cetegen, B.M., Kasper, K.D.: Experiments on the oscillatory behaviour of buoyant plumes of helium

and helium-air mixtures. Phys. Fluids8, 2974–2984 (1996).


2

AEDSS-2014-102

Velocity and Density field of compressible vortex rings

C. L. Dora and Debopam.Das†

, † Department of Aerospace Engineering, Indian Institute of Technology Kanpur-208016, India Corresponding author‟s email: [email protected] Corresponding author‟s Ph. No.: +91-9956817221

Abstract

The exhaust flow of rocket nozzle start-up, transient supersonic circular jets, volcanic outbursts, and Muzzle blasts are

characterized by shockwaves and compressible vortex rings. In general, a circular vortex ring is produced when an impulsive flow

exits from a nozzle or an orifice. On the other hand, compressible vortex rings are usually generated at the open end of a shock

tube. The present study aims at exploring the flow and density field of compressible vortex rings by experimental and

computational methods with particular emphasis on counter rotating vortex rings (CRVR) formation and their dynamics. In the

earlier studies, the embedded shock strength has been asserted as the primary cause for formation of CRVR. However, in the

present study particle image velocimetry measurements and numerical simulations show that CRVR does not form in the

absence of Mach disk in sonic under-expanded jet behind the primary vortex ring. The Kelvin-Helmholtz type shear flow instability

of the slipstream originated from the triple point of the Mach disk, and subsequent eddy-pairing, as observed by Rikanati et al.[1]

in shock-wave Mach-reflection, is found to be responsible for CRVR formation (see Fig.1). The growth rate of the slipstream in

the present problem follows the model proposed by them. Additionally, the parameters influencing the formation of CRVR as well

as its dynamics are investigated. It is found that the strength of the Mach disk and its duration of persistence results in an exit

impulse that determines the number of CRVRs. Furthermore, the density field of compressible vortex rings is evaluated by

employing Background oriented schlieren Technique (BOS). It has been observed that the core of these compressible vortex

rings are characterized by steep density gradient following a gaussian profile. However, the tangential velocity within the core

follows a linear relation similar to that of a incompressible vortex ring.

Figure.1: Smoke flow visualization showing vortex ring, trailing under-expanded jet, and slipstream at the open-end of a shock tube.

References: [1].Rikanati, A and Sadot, O and Ben-Dor, G and Shvarts, D and Kuribayashi, T and Takayama, K. “Shock-wave Mach-reflection slip-stream instability: a secondary small-scale turbulent mixing phenomenon “ Physical review letters, 2006, 96, 17450

mailto:[email protected]


3

AEDSS-2014-103

Experimental Aerodynamics of a Butterfly mimicking Flapping Wings at Low

Reynolds Number

Saurav K Ghosh

a and Debopam Das

Department of Aerospace Engineering, Indian Institute of Technology Kanpur, Kanpur


The unsteady aerodynamics of a butterfly mimicking flapping wings under zero and non-zero advance

ratio is investigated experimentally in the Re range of 2000-1000. The present study encompasses two

experimental aspects of unsteady aerodynamics of the flapping wing regime; the force measurement [1]

and the PIV measurement [2]. By the above mentioned approach, the significance of LEV (leading edge

vortex) has been emphasized in the unsteady force generation. The role of the chord-wise flexibility is

substantially studied. The dependence on the kinematics of the wing is also noted by comparing two

flapping models, one with simple flapping and one with lead lag along with flapping motions. The impact

of clap and fling on particularly the lift force is observed in an extensive way. PIV experiments serve both

as a tool of flow visualization as well as a quantitative judgment for studying the nature of flow over a

flapping wing. Furthermore, the observations with the PIV analysis are compared with the force results to

explain the nature of the forces and their patterns. The flow is observed to be more attached and stable for

the lead lag kinematics of the wing. However, the ejection process noted clearly in the normal flapping

isn’t that visible with the lead lag motion. The butterfly shaped wing shows better lift and thrust

generation than a rectangular wing of the same span and same aspect ratio (AR).

Figure 1: (a) Schematic view of the model (the dimensions are in mm) (b) wing mounted on a linkage (c)

comparison of lift forces in a butterfly shaped wing and a rectangular wing of the same span and same AR

References:

3. Abhijit Banerjee, Saurav K. Ghosh, and Debopam Das (2011) , “Aerodynamics of flapping wing at low

reynolds numbers: force measurement and flow visualization,” ISRN Mechanical Engineering, vol. 2011,

Article ID 162687, 8 pages, 2011. doi:10.5402/2011/162687

4. Ghosh, S.K., Dora, C., and Das, D. (2012). ”Unsteady Wake Characteristics of a Flapping Wing through 3D

TR-PIV.” J. Aerosp. Eng. 25, SPECIAL SECTION: Intelligent Unmanned Systems, 547–558.

(b) (c)

(a)


4

AEDSS-2014-104

Viscous flow in a twin intake at supersonic speeds

V M Krushnarao Kotteda and Sanjay Mittal

Department of Aerospace Engineering, IIT Kanpur, Kanpur, Uttar Pradesh-208016 [email protected], [email protected]

Air Intakes are one of the major components of any aircraft engine. Air intakes are essentially a fluid flow

duct whose task is to ensure the engine functions properly to generate thrust. The intake must be designed to provide

the appropriate amount of airflow at low speed with relatively low distortion required by the engine. The design of

the intakes generally aims to provide the shortest possible duct length for a given cross-sectional shape at inlet and

outlet in order to minimize the friction losses and flow distortion at the exit by avoiding separation. The shape of the

intake is generally taken as curved with increasing area from inlet to exit. The curved intakes are generally known as

Y-intake, twin and bifurcated intakes. These are referred to a pair of intakes in the wing roots or on the two sides of a

fuselage, feeding a single engine via a common section of duct or a plenum chamber. With air intakes on the

fuselage, notably the Pitot type, the normal shock positioned at the entrance interacts with the fuselage boundary

layer. This is generally very much attenuated by the presence of a splitter plane just upstream. It should be noted that

this interaction, while limited, may induce a slight compression by oblique shock, which is more favorable than the

nominal normal shock. This interaction favors the starting of air intakes with mixed compression. Y-intakes are

extensively used for engines in fighter aircrafts. These ducts normally operate under steady and symmetric

conditions. Air flows in two intakes join in a common duct are subjected to air-flow instability [1-3] at low mass

flows. A particular type of instability, which is characterized by fluctuations of the quantity of the flow in each duct

and which usually results in reversal of flow in one of the ducts as the mass flow is reduced further. Instability in the

intake adversely affects the mass flow entering the engine and may lead to combustion instability, engine surge and

flame out. It can also lead to deterioration of the performance of propulsion system, thus causing catastrophic loss in

thrust.

The viscous flow in a Y-shaped intake in two dimensions has been studied via a stabilized finite element method.

The free stream Mach number of the flow entering the intake is 1.5 and the Reynolds number, based on the height of

the intake, is 105. Computations have been carried out for various sideslip angles. First, the flow is computed for the

situation when the outflow is supersonic and no condition on back-pressure, pb, exists. The flow is utilized to carry

out computations for various values of back pressure ratio, pb/pi. The strength of the shock increases as pb/pi

increases. It moves upstream of the merger section as pb/pi increases. At the critical operation it is positioned at/near

the leading edge of the cowl. The instability occurs when pb/pi increases beyond critical pb/pi. The mass flow rate at

the merger section decreases when pb/pi increases. The total pressure recovery increases with pb/pi up to the critical

pb/pi and decreases for pb/pi beyond critical pb/pi. As the sideslip angle, β increases the critical pb/pi decreases. The

instability range or flow asymmetry range increases with β. Steady flow range in the intake decreases as β increases.

References

[1] Martin NJ, Holzhauser CA, 1950, Analysis of factors influencing the stability characteristics of symmetrical twin-intake air-induction systems, NACA TN–2049.

[2] Seddon J, Trebble W, 1955, Model test on the asymmetry of airflow occurring in twin intake systems at subsonic speeds, Aeronautical Research Council Reports and Memoranda–2910.

[3] Anderson WE, Perkins EW, 1959, Effects of Unsymmetrical Air-flow Characteristics of Twin-intake Air-induction Systems on Airplane Static Stability at Supersonic Speeds, NACA TM X–94.


5

AEDSS-2014-105

EXPERIMENTAL INVESTIGATION OF ASYMMETRIC PITCHING OSCILLATIONS

OF A SYMMETRIC AIRFOIL

Anshul Khandelwal1,a

, K. Poddar2 and D. Das

3,

1PhD student, IIT Kanpur 2Professor, IIT Kanpur

3Associate Professor, IIT Kanpur


The aerodynamic characteristics of an airfoil in unsteady motion have many important

implications, and this has been the subject of study of numerous researchers. The present study is

directed towards dynamic effect of asymmetric, unsteady, pitching oscillations of an airfoil on

the flow development and resulting aerodynamic loads on the airfoil at relatively low Reynolds

number which may find application in MAVs and vertical axis wind turbines. Investigation

highlights the results for the conditions when the airfoil is not stalled and dynamically stalled. In

this paper, the effect of asymmetry parameter S, defined as the fraction of the time-period of one

cycle required to reach the minimum amplitude starting from the maximum, on unsteady flow

development on NACA 0012 airfoil at a Reynolds number 8.5*104 has been investigated at

different mean angles and reduced frequencies through unsteady pressure measurements on the

mid-span of airfoil. Pressure data has been ensemble-averaged over 500 cycles to remove the

randomness due to cycle-to-cycle variations in flow and helps to quantify the average load

distribution at different phase angles.

It has been found that introducing asymmetry in oscillation at a fixed value of reduced

frequency has profound effect on unsteady flow development. It significantly affects the

formation, strength and detachment of dynamic stall vortex and consequently the pressure

distribution on airfoil surface, integrated loads and the shape & size of hysteresis loops. The

effect is seen to be most dramatic in the deep dynamic stall regime. Figures 1 to 3 show the effect

of asymmetry on the hysteresis loops at mean angle 10° and reduced frequency 0.1. Higher the

asymmetry, more is the deviation and resulting difference in behavior from symmetric

oscillations; in particular for the case of oscillations with faster upstroke cycle. The reason for

change in shape of hysteresis loops has been discussed.


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AEDSS-2014-105

Fig 1(a): κ = 0.1, S =0.5

Fig 1(b): κ = 0.1, S =0.67

Fig 1(c): κ = 0.1, S =0.33

Figure 1: Effect of S on phase-averaged lift & moment coefficients on NACA 0012 airfoil


7

AEDSS-2014-106

Velocity Dynamics of Suddenly Blocked Oscillatory Channel Flow Avinash Nayak1,*, Ashok Kannaiyan1 and Debopam Das1

1 Department of Aerospace Engineering, IIT Kanpur, India. 208016. * Email: [email protected] * Ph. No.: +91-9005813019

ABSTRACT

In internal flow systems, it is frequently required to rest the flow suddenly. Due to the rapid deceleration of the flow

which may be periodic or non-periodic, the flow characteristics such as velocity, pressure and shear stress also

change substantially during this short interval. So, an accurate quantitative knowledge of flow dynamics in this

phase of flow is very important for design and control of internal flow systems and physiological flows.

Much study has been done on the suddenly blocked non-periodic laminar fully developed poiseuille flow. The work

of Weinbaum [1] along with studies of Das [2] and Scandura [3] illustrates the flow dynamics when a fully

developed laminar flow is suddenly blocked by application of end wall or a valve.

But not much work is available in literature for blockage of unsteady flows like an oscillating flow. So, in the current

study, an exact solution to the flow dynamics of laminar pulsatile flow in a channel which encounters a sudden

blockage is attempted. The solution is obtained using Das-Arakeri approach [2], which is comparatively simple but

powerful. Velocity profiles, pressure gradient and shear distribution during decay time are obtained. The solution is

valid for various parameters like Womersley number and the phase of cycle at which the flow is blocked. This

knowledge has an important role in engineering and physiological phenomena, such as an instance of closing,

restarting of fluid control systems, and blood flow in arteries.

Figure 1: Decay of velocity profile with time for initial

phase = 90o; o: t = 0.0001; : t = 0.1; +: t = 1; ×: t = 2.

Figure 2: Pressure gradient variation with time; o: W =

4; : W = 10.

References

[1] Weinbaum, S and Parker, K H, 1975, The laminar decay of suddenly blocked channel and pipe flows, Journal of Fluid

Mechanics, 69(04), 729–752.

[2] Das, D and Arakeri, J H, Unsteady laminar duct flow with a given volume flow rate variation, Journal of applied

mechanics, 67(2), 274–281.

[3] Scandura, P, 2003, Two-dimensional perturbations in a suddenly blocked channel flow, European Journal of Mechanics-

B/Fluids, 22(4), 317–329.


8

AEDSS-2014-107

Numerical Investigation of Clap and Fling Motion Using Immersed Boundary Lattice Boltzmann Method

Pradeep Kumar S, Ashoke De*and Debopam Das*

*Department of Aerospace Engineering, IIT Kanpur, Kanpur - 208016, INDIA Corresponding author‟s email: [email protected]

Corresponding author‟s Ph. No.: +91-8960419188

Extended Abstract

Clap & Fling, a high lift producing mechanism in insects flight is simulated using an Immersed Boundary Lattice Boltzmann method. Effects of advance ratio and frequency on flow field at different Reynolds number have been studied. The numerical results agree well with the existing experimental & numerical data.

1. Introduction

This paper aims at simulating the clap and fling motion first proposed by Weis Fogh in

1973 using immersed boundary Lattice Boltzmann method. LBM has emerged as a powerful

numerical technique since past decade for simulating fluid flows. In this simulation, LBM

with single relaxation BGK model is used to solve Boltzmann transport equation for particle

distribution function from which the macroscopic properties of the fluid can be recovered by

evaluating the moments of the distribution function. The fact that both IBM & LBM are

applied around the Cartesian mesh, makes this combination both attractive and possible.

2. Formulation

In the current simulation, two rigid elliptic airfoils are made to clap and fling using an

implicit velocity correction method introduced by Wu Shu et al.[1] to simulate. A D2Q9

Boltzmann model with forcing scheme as introduced by Guo et al [3] is used.

The Physical velocity at a fluid node is assumed to be sum of intermediate velocity (u*)

(obtained from the moment of distribution function) & correction velocity(due to the discrete

lattice effect).

where &

Unlike conventional IBM methods, like direct forcing or momentum exchange method, the

force density here is considered as unknown & is determined in such a way that the boundary

(Lagrangian) velocities obtained from interpolated corrected velocity field (Eulerian) satisfies

the no slip condition. A dirac delta function interpolation is implemented as proposed by

Peskin.


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AEDSS-2014-107

These boundary velocities are used to calculate the velocity correction at the Eulerian

points using a spreading function which then will be introduced in Boltzmann equation which

will make the fluid aware of the boundary through the forcing term.

3. Result

The code is validated with the benchmark problems like flow over a square and circular

cylinder. The fig.1 shows the vorticity contours of four time instances in one complete clap &

fling cycle. The fig 2 shows the corresponding time history of lift and drag coefficients

Fig 1: Vorticity Contour during 2nd Cycle for Re =100,f=0.05Hz

a) b) Fig:2 Time histories of a) Drag coefficient b) Lift Coefficient at Re = 100, f=0.05 Hz.

4. Conclusion

The Implicit IB LBM method is able to capture the flow physics like wake capture,

dynamic stall, wing - wing interaction, LEV which makes this a robust method to investigate

insect flight phenomenon. This work in future can be extended by introducing flexibility into

the wing structure.

5. References

1. J. Wu, C. Shu: Implicit velocity correction-based immersed boundary-lattice Boltzmann method and its applications. J.

Comput. Physics 228(6): 1963-1979 (2009)

2. Charles S. Peskin (2002). The immersed boundary method. Acta Numerica, 11, pp 479-517.

3. Weis-Fogh, T. (1973). Quick estimates of flight fitness in hovering animals, including novel mechanisms for lift

production. J. Exp. Biol. 59, 169-230.

4. Z. J. Wang, “Two-dimensional mechanism of hovering,” Phys. Rev. Lett. 85, 2216 (2000).

5. S. K. Mishra, A. De (2013), "Coupling of reaction and hydrodynamics around a reacting block modeled by Lattice

Boltzman Method (LBM)", Computers and Fluids, Vol. 71, pp. 91-97

http://www.informatik.uni-trier.de/~ley/pers/hd/w/Wu:J=.html

http://www.informatik.uni-trier.de/~ley/db/journals/jcphy/jcphy228.html#WuS09

http://www.informatik.uni-trier.de/~ley/db/journals/jcphy/jcphy228.html#WuS09

http://link.aip.org/link/?&l_creator=getabs-normal1&l_dir=FWD&l_rel=CITES&from_key=PHFLE6000020000008087101000001&from_keyType=CVIPS&from_loc=AIP&to_j=PRLTAO&to_v=85&to_p=2216&to_loc=APS&to_url=http%3A%2F%2Flink.aps.org%2Fabstract%2FPRL%2Fv85%2Fp2216


10

AEDSS-2014-108

NORMAL INTERACTION OF A COMPRESSIBLE VORTEX RING

AND WALL

D.Saravanan, Debopam Das†

*†Department of Aerospace Engineering, IIT Kanpur, Kanpur, UP, 208016, Corresponding author‟s email:[email protected] Corresponding author‟s Ph. No.: +91-9889288665

Extended Abstract

The normal-interaction of a compressible vortex ring with a plane surface is studied both experimentally and numerically. Isolated vortex rings are generated using an open-ended shock tube in both experiments and simulation for shock Mach number range of 1.31≤ M ≥1.88 for different wall distance X/Dri= 1.85 and 2.91 where Dri is the initial ring diameter. The trajectory of the vortex ring from the initial to the point of separation of the vortex and also after separation forms a fundamental study to determine some common phenomena for all Mach numbers. The shock-vortex interactions and the unsteady separation of the vortex ring during wall interaction are quite complex and thus plays a major role in altering both the geometric parameters (ring and core diameters) and the kinematic characteristics (Circulation) of the vortex and hence these are quantified to know the effects on the vortex.

1. Introduction

Shock waves and vortices are present everywhere on vehicles traveling at supersonic speeds and hence their interactions also occur in many situations of practical interest like supersonic aircraft and missiles. This shock associated noise is an important factor in design of advanced jet engines. Other fields which include these interactions are the case of helicopter blades operating at super critical speeds and fuel-air mixing in the combustion of a supersonic combustion ramjet. This normal shock-vortex interaction can be triggered experimentally by bursting the of a diaphragm in a shock tube resulting with a shock-induced vortex ring and concentrated vortices at the open ended shock tube after translating and interacting with the shock finally move onto a wall. The entire sequence can be described as (i) vortex evolution, (ii) the interactions are accordingly shock-shock (embedded shock), shock and the vortex interactions depending on the shock Mach number, (iiI) the vortex stretching (a wall effect), (iv) finally their interactions on the wall (impingement) (v) the emergence of the separation bubble usually called secondary wall vortex and (vi) finally the lift-Off. The geometric parameters ring diameter and core diameter and the kinematic characteristics are calculated as given by Arakeri et al [1]. Murugan and Das [2] also

obtained the entire sequence using smoke flow visualisations using driver section lengths of Ldrvr/D=1.77 for M=1.31 and 1.55 and Ldrvr/D=4.9 for 1.7 and 1.88 at X/D = 4.7 with the same driven section length used in this study.

2. Experimental Setup

Experiments are performed using an open ended shock tube, which is shown schematically in figure 1. The inner and outer diameters of the shock tube are 64mm and 100mm, respectively. Experiments are performed with a constant driven section length (Ldrvn/D= 18.75, where D is the inner Diameter of the shock tube.) for four different Mach number, M =1.31, 1.34, 1.55 and 1.88. Mylar sheets are used as diaphragms and they are ruptured using a thin (0.5mm) electrically heated copper wire. With Helium as driver section gas, the diaphragm pressure ratios for the respective shock Mach numbers mentioned above are approximately 3.0, 2.7, 5.5 and 11.5, respectively. Particle Image Velocimetry (PIV) technique has been employed for measuring



11

AEDSS-2014-108

the velocity field of the flow coming out from the shock tube, with a double-pulsed Nd-YAG laser (100mJ/pulse, 50Hz, thickness of the sheet ' 1mm) as the illumination source and a 12-bit CCD camera of 4MP (Mega Pixel - 2048 pixels X 2048 pixels) resolution to capture the image data, with a frame rate of 4 Hz in double frame mode. Adaptive cross correlation PIV algorithm was used to analyze the captured flow images using an interrogation area of 32 pixels X 32 pixels and 50% overlap in the final step of the processing.

3. Result The velocity field in the figure 1 (a) and (b) shows the interaction of the reflected shock with the vortex before and during the impact. The horizontal extraction of velocities U, V and Length (Total Velocity) along the axial direction at the center of the vortex will show fluctuations in the U velocity (which is not shown here). This will show that the interaction of the reflected shock with the vortex will affect only the quantity U, but not the remaining quantities. Thus it also affects the core-diameter and the circulation and hence the vorticity. Hence the vortex ring suffers a shrink in core-diameter, a loss in translational velocity, and circulation. After the interaction the vortex ring again regains and these quantities begin to increase due to stretching and it again falls during the formation of separation bubble.

Figure 1: Particle Image velocimetry of the vortex ring (a) before (Left) and (b) during (Right) the reflected shock interaction for Ldrvn/D= 1.02 for X/Dri=1.85, where Dri is the initial ring diameter.

4. Conclusion The trajectory of the vortex ring is a fundamental measure of the variation of the geometric parameters such as ring and core-diameter which have implications on kinematic characteristics such as circulation. The interaction of the reflected shock with the vortex ring mainly affects the axial velocity (U).

5. References

[1] J. H. Arakeri, D. Das, A. Krothapalli, and L. Lourenco, 2004, Compressible vortex ring: a PIV study, Physics of Fluids, 16(4), 1008–1019. [2] Murugan Thangadurai,Debopam Das, 2012,Experimental Study on a compressible vortex ring in collision with a wall,J. Vis,15, 321-332

(a) (b)


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AEDSS-2014-109

Free vibrations of a cylinder beyond the laminar regime

Navroseand Sanjay Mittal

Computational Fluid Dynamics lab

Department of Aerospace Engineering, Indian Institute of Technology Kanpur, Kanpur, Uttar Pradesh 208 016

[email protected]

Extended Abstract

Unsteady aerodynamic forces arising from vortex shedding from a bluff body may lead to its vibration. There is a

strong interdependence between the motion of the body and the wake structure. Free vibrations are often

associated with resonance/synchronization/lock-in, wherein the frequency of oscillation of the structure matches

the frequency of the periodic wake vortex mode (Khalak and Williamson 1999). More often than not, the

synchronization regime is associated with hysteretic and /or intermittent behavior. There have been numerous

efforts in the past to study the lock-in phenomenon associated with a vibrating circular cylinder. For a

comprehensive review the interested reader is referred to review articles by Bearman (1984, 2011), Williamson

and Govardhan (2004), Sarpakaya (2004) and Wu et. al. (2012).

An important parameter that affects the response of the cylinder in the synchronization regime is the Reynolds

number. In the laminar regime, synchronization is associated with two branches of response: initial and lower

(Prasanth and Mittal 2008). At large Reynolds numbers (Re ~ 10, 000) experimental studies indicate presence of

three response branches: initial, upper and lower (Khalak and Williamson 1999, Govardhan and Williamson 2000).

The two kind of response curves are shown in Figure 1. Despite significant amount of work in the two flow

regimes, the exact reason for the difference in the response between the two regimes is unexplored.

Figure 1: Variation of amplitude with reduced velocity. The reduced velocity is defined as the inverse of non-

dimensional natural frequency of the spring mass system. The data for laminar regime is taken from Prasanth

and Mittal (2008), while that for high Re is borrowed from Govardhan and Williamson (2000).


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AEDSS-2014-109

In this study, we first present results from 3D computations carried out on a circular cylinder that is allowed to

oscillate for Re = 1000. The aspect ratio of the cylinder is AR = 4.0, and non- dimensional mass in m* = 10.0. The

motion of the cylinder along the cartesian axes is governed by two degree of freedom spring mass system. A

stabilized space- time formulation is utilized to solve the flow equations. Towards the end we will present some

preliminary results from on-going computations for 3000 < Re < 10,000.

References:

1. Khalak, A., Williamson, C. H. K., 1999. Motions, forces and mode transitions in vortex- induced vibrations at

low mass- damping. Journal of fluids and Structures 13, 813-851 .

2. Bearman, P. W., 1984. Vortex shedding from oscillating bluff bodies. Annual Review of Fluid Mechanics 16,

195-222 .

3. Bearman, P. W., 2011. Circular cylinder wakes and vortex- induced vibrations. Journal of Fluids and

Structures 27, 648-658 .

4. Williamson, C. H. K. and Govardhan, R. N., 2004. Vortex- induced vibrations. Annual Review of Fluid

Mechanics 36, 413-455.

5. Govardhan, R. N. and Williamson, C. H. K., 2006. Modes of vortex formation and frequency response of a

freely vibrating cylinder. Journal of Fluid Mechanics 420, 85-130.

6. Prasanth, T. K. and Mittal, S., 2008. Vortex- induced vibrations of a circular cylinder at low Reynolds numbers.

Journal of Fluid Mechanics 594,463-491.


14

AEDSS-2014-110

Analysis of Aerodynamic Forces and Inflight Measurements Of An

Ornithopter

Joydeep Bhowmik, Nidhish Raj† and Debopam Das

*Phd Student, IIT Kanpur, City, UP, 208016.

† Research Associate, IIT Kanpur, City, UP, 208016

Associate professor, IIT Kanpur, City, UP, 208016

Corresponding author’s email: [email protected]

Corresponding author’s Ph. No.: +91-9793430022

Extended Abstract

An ornithopter is mechanical replica of a bird which flies by flapping its wings. Lift and thrust is solely

produced from the flapping of the wings but unlike a fixed wing aircraft, the aerodynamic forces and

moments vary with time. The overall effect of these time varying forces affects the kinematics of the vehicle

and essential to understand the same to develop such a vehicle into bio-mimicking stealth surveillance

UAV carrying a camera. Theoretical analysis has been done and compared with some existing methods for

a given wing with a given flapping kinematics. Force measurements on an ornithopter are also carried out

in a wind tunnel but since the model is rigidly fixed to the force transducer, the combined effect of the

aerodynamic forces due to the movement of the vehicle while flight is obscured.

This work presents an experimental approach to determine the aerodynamic forces that act on an

ornithopter in real flying conditions.

In the previous work a theoretical model based on unsteady lifting line theory has been developed and

several ornithopters have been successfully built and tested. Wind tunnel test reveals dominance of the

highly unsteady nature of aerodynamic forces which are not assumed in this model. The model has also

been modified to get the lift and thrust at different position of the wing to get more accurate estimate of the

aerodynamic force during the flapping cycle.

UAV’s now a day have developed a lot, It is not difficult to see an autonomous aircraft of a rotary wing

vehicle accomplishing complicated tasks without human control. But ornithopters on the other hand are at

the developing stage as the unsteady nature of aerodynamic forces offers a challenge in anticipating the

same. The present study focuses mainly on the study of aerodynamic forces and moments that are

generated on an ornithopter and their effect on the flight of the vehicle in a steady level cruise flight.

Figure 3 shows the result of the cyclic variation of aerodynamic forces for a wing flapping at 4 Hz estimated

using an unsteady aerodynamic model using modified Theodorsen’s lift deficiency function C(k).


15

Figure 3 (a) Variation of Thrust with non dimensionalised cycle time (b) Variation of Lift with non dimensionalised cycle

time.

.

0 0.2 0.4 0.6 0.8 1-0.1

-0.05

0

0.05

0.1T

hru

st

(N)

0 0.2 0.4 0.6 0.8 1-40

-20

0

20

40

Win

g p

ositio

n (

degre

es)

cycle time (t/T)

Thrust

wing position

0 0.2 0.4 0.6 0.8 1-1.5

-1

-0.5

0

0.5

1

1.5

Lift

(N)

0 0.2 0.4 0.6 0.8 1-30

-20

-10

0

10

20

30

Win

g p

ositio

n (

degre

es)

cycle time (t/T)

Lift

wing position


16

AEDSS-2014-201

Emission Measurements from Bench Scale Aircraft Combustor Rig

V. Pandey, A. Kushari†, A. De‡, J.S. Dinesh Kumar˚, A. Raut††

* Doctoral Student, Aerospace Engineering, IIT Kanpur, UP, 208016.

† Associate Professor, Aerospace Engineering, IIT Kanpur, UP, 208016. ‡Assistant Professor, Aerospace Engineering, IIT Kanpur, UP 208016 ˚ Doctoral Student, Aerospace Engineering, IIT Kanpur, UP, 208016.

†† Research Associate, Aerospace Engineering, IIT Kanpur, UP, 208016. Corresponding author‟s email: [email protected]


Extended Abstract

Present day global scenario for crude oil production indicates a general overall decline. On the other hand, there are trends for steady increase in global consumption of crude oil. This would cause a widening of the demand-supply gap in the foreseeable future. A similar trend is observed for the prices of aircraft engine fuels as well since aviation turbine fuels (ATF) are derived as a middle distillate from crude oil. There is therefore a global imperative by most major airlines to find a substitute to ATF that will be cheaper and also have a small environmental footprint. Extensive experimentation goes into this process for the purpose of emission evaluation.

1. Introduction

Space and weight are primary constraints in aircraft engine design and much effort is made to reduce combustion chamber volume by using energy dense liquid fuels that have high specific energy (of the order of 45 MJ/kg). In addition, storage and transport of liquid fuels are safer and easier compared to gaseous fuels [1]. Hence liquid fuels are indispensable for aircraft applications. ATF is used for aircraft engines and it a middle distillate of crude oil. Due to the rapid technological changes in the 20th century and concurrent rapid depletion of fossil fuels, we have crossed the „peak‟ of production curve for crude oil and by implication, its derivatives. The „peak oil‟ theory was propounded by various scholars in the first half of the 20th century and it has now found wide credence and acceptability. The theory suggested that the peak of world oil production would occur in the first decade of the 21st century. The International Energy Agency reports that world peak for crude oil occurred in 2006. Reports also predict a continuous, long term and global trend of increase in consumption of these fuels, especially by BRIC economies, thereby widening the demand-supply gap. At present, there is a global, concerted effort by the aviation sector to develop and test newer fuels with comparable energy densities and properties so that ready substitutes to ATF are available. Furthermore, the burning and emission characteristics of the newer fuels should be the same as the fuel they would replace. Emissions from aircraft engines in particular must conform to stringent cut-offs. The environmental impact from emissions has become well known in recent years. In addition to the effect on the global environment, there is a local penalty associated with emissions. Different nations have their own emission standards, some more strict than others. An aircraft engine must conform to the severest of prevalent emission norms or else a penalty is levied for excess. Emission characteristics are inexorably linked to combustion performance and gas turbine/combustor design.

2. Formulation

For the purpose of testing ATF for emissions under varying flight conditions, a bench test rig has been developed at Combustion and Flame Dynamics Laboratory at IIT Kanpur. Bench tests are necessary prior to full scale, in flight testing. For this purpose, a can-combustor [2] (single can of a can-annular combustor) is used for testing. The bench test rig comprises of the following: Air Compressor of 1100 CFM, 300 psi capacity at 2200 rpm, a


17

AEDSS-2014-201

Vortex Flow Meter, Orifice Flow Meter for air flow measurement, a 200 kW Electric Furnace designed to withstand 20 bar pressure at 1300 K (melting point of steel), closed loop controlled inlet and exit valves, Coriolis Flow Meter for fuel flow measurement, a high pressure jet pump for after cooling of exhaust prior to exit, an Emission Analyser with modules for the measurement of NOx, CO, CO2, O2, Total Hydro-Carbons (THC) and lastly a comprehensive Data Acquisition System by LabView® along with pressure and temperature acquisition devices.

3. Result Experiments were carried out at various test conditions, one of which is mentioned below in Table1. The NOx ppm obtained is also depicted in the Figure 1. Experimental results match with computational results carried out earlier.

P3(bar) dP dP (%) air

flow

rate

(kg/sec)

FAR fuel flow

rate

(kg/sec)

AC

Drive

Freq

(hz)

Pump

RPM

T3

(deg

C)

T4

(deg

C)

1.34 0.076 0.056 0.059 0.031 0.00186 17 570 194 593

Table 1. Test Condition

Figure 1. NOx ppm from test condition

4. Conclusion For the purpose of measuring emission characteristics using a gas turbine combustor bench rig, experiments were done and emissions measured. NOx measurements, pressure drop across combustor and T4 were measured and these correspond closely with computational results carried out in previous study.

5. Acknowledgement

Support from Pratt and Whitney is acknowledged for this work.

6. References

[1] Weinberg, F., 1974, The first half million years of combustion research and today‟s burning problems, Plenary Lecture, 15th (Int.) Symposium on Combustion, Tokyo, Japan. [2] Sampath, P, Shum, F, 1985, Combustion Performance of Hydrogen in a small Gas Turbine Combustor, Int. J. Hydrogen Energy, 10:12, pp. 829-837.


18

AEDSS-2014-202

Large Eddy Simulation of Swirling Non Reactive Flow

Suresh Balaji V, Ashoke De†

* M.Tech Student, IIT Kanpur, Kanpur, 208016. † Assistant Professor, IIT Kanpur, Kanpur, 208016. Corresponding author‟s email:[email protected].

Corresponding author‟s Ph. No.: +91-512-259-7863

Extended Abstract

In the present work, Large Eddy Simulation technique is used to study the unsteady flow structures of turbulent swirling non reacting flow. Two non reacting flows from Sydney swirl flame database is used for validation purpose. The temporal evolution of the flow is studied using the pressure iso-surface. Two reverse flow zones with a collar-like structure between them is observed in low swirl case. In high swirl case, a single long reverse flow region is observed. Q-criterion technique is used to study the complex vortex structures.

1. Introduction

Efficient use of energy and reduction of emissions are the most essential design parameters for engineering appliances, especially in propulsive devices. This gives rise to the need for constant progress in the research areas of combustion to meet the stringent emission regulations and environment requirements. Combustion in swirling flows is one of the most widely used techniques as it shows a significant enhancement in stability of the flame with the help of recirculation zones. Besides enhancing stability, better mixing is achieved which gives rise to increase in combustion efficiency thereby reducing the pollutant emissions. In addition to that, it also reduces the combustion length by increasing the rate of mixing. Hence the modeling of swirling flames is always been an emphasized task of interest in the combustion community. To have better understanding of these kind of flames, modeling of non reactive flows is considered to be very vital. In this present study, Large Eddy Simulation (LES) technique is used to study the unsteady flow structures and the vortex breakdown processes involved in two isothermal cases (N29S054 & N16S159) from Sydney swirl flame database[1].

2. Formulation

In Large Eddy simulation technique, the large scale energy containing structures are resolved numerically, while the small scale unresolved structures are modeled. Applying the box filter, the filtered continuity and momentum equations for the large scale structures are as follows.

0i

i

u

x

(1)

( ) (2 ) ( )1i j ij iji

j i j j

u u Su P

t x x x x

(2)

where the strain tensor, 1

2

jiij

j i

uuS

x x

The last term in equation (2) is known as Subgrid scale(SGS) tensor. It represents the contribution of subgrid

scale to the momentum equation. For closing the equation, the term, ij i j i ju u u u has to be modeled.

Smagorinsky Eddy Viscosity model[2] is used to model SGS term. Here the eddy viscosity, sgs is a function of the

filter size and strain rate.



19

AEDSS-2014-202

2

sgs sC S (3)

where sC is a Smagorinsky model parameter and

1

22 ij ijS S S . The dynamic procedure [3] is used to

calculate the model parameter sC .

3. Result The simulations are done using commercial code ANSYS FLUENT- 15.0. The numerical predictions are compared with experimental results. The predictions show overall good agreement with the experimental measurements. Figure 1 shows the evolution of unsteady flow structures for high swirl case using pressure iso-surface and axial velocity contours. The Q criterion technique is also used to study the complex turbulent structures involved in the flow.

t=0.045 sec t=0.048 sec t=0.066 sec

Figure 1. Spatial-Temporal evolution from Isosurface of Pressure (-462 Pa) and contour axial velocity from 0.045 sec to 0.066 sec for N16S159

4. Conclusion Large Eddy simulation technique has been successfully applied to study the unsteady flow structures of two swirling non reacting jets from Sydney swirl database. The numerical predictions have been compared with the experimental measurements. The pressure iso-surface in combination with instantaneous axial velocity is used to study the temporal evolution of the flow. The vortex break down process involved is studied using Q-Criterion technique. A commercial CFD code is successfully applied in studying the complex flow structures using Large Eddy Simulation technique.

5. References

[1] http://sydney.edu.au/engineering/aeromech/thermofluids/ swirl.htm [2] Smagorinsky, J, 1963, General circulation experiments with the primitive equations, M. Weather Review, Vol.91, pp.99-164 [3] Piomelli, U. and Liu, J, 1995, Large eddy simulation of channel flows using a localized dynamic model, Phy. Fluids, Vol. 7, pp. 839-848.


20

AEDSS-2014-203

Effect of transverse periodic loading on an airfoil in a cascade

M. C. Keerthi, Abhijit Kushari†

*Graduate Student †Associate Professor

Department of Aerospace Engineering, IIT Kanpur, UP, India 208016. Corresponding author‟s email: [email protected]

Corresponding author‟s Ph. No.: +91-512 259 7851

Extended Abstract

Aircraft gas turbine engines are being subjected to increasingly severe requirements in terms of thrust to weight ratio. The blades of axial turbomachines are therefore subjected to higher loading while the blade thickness is getting smaller. As a result, strong fluid-structure interaction results between the unsteady air stream and the vibrating blades. Due to the effect of adjacent blades, the aeroelastic phenomena in turbomachines establishes in a manner different from an isolated aircraft wing aeroelasticity. The present study is a preliminary CFD study of damping performance for a compressor cascade over a range of relevant parameters.

1. Introduction

Present-day aircraft gas turbine designs are tending towards maximize the performance at lower costs, without compromising on safety. In the context of axial turbomachines, this entails having thinner blades while being subjected to the same or higher aerodynamic loading. Due to the high relative velocities encountered, this can lead to a self-starting and sustaining blade vibratory phenomena called flutter. The prevailing unsteadiness in the air stream, due to upstream rotor wakes or gusts, can lead to blades vibrating at their natural frequencies, known as forced response. Both flutter and forced response cause high-cycle fatigue leading to catastrophic blade failures. The current predictive tools for aeroelasticity in turbomachines are not adequate possibly resulting in a costly redesign of the blading in the later stage of engine design.

Some of the critical parameters in turbomachine aeroelasticity are the interblade phase angle and the reduced frequency. The present study is a fundamental CFD study on the effect of interblade phase angle on the damping characteristics of a reference stationary airfoil.

2. Formulation

In an actual turbomachine, the blades vibrate with a non-zero phase difference with each other. This is enforced by the nodal diameter of the rotor disc, resulting in a constant phase difference between any two adjacent blades. The phase difference between the forces acting on a vibrating blade and the motion of the blade is a critical parameter that decides whether the amplitude of the blade vibration can grow in time. The net work done on the structure is a function of this parameter, and its sign is a direct indicator of instability [1]. In the present study, only the effect of the oscillating neighboring blades on a stationary airfoil is studied.

The cascade under study consists of five airfoils with a spacing of 60.5 mm. The airfoil is a standard test configuration 1, recommended for low-speed compressor aeroelastic studies Bölcs and Fransson [2]. The camber angle of the airfoil is 10° and the chord is 121.7 mm. The stagger of the blades is set at 0°. In the present study, the second and fourth blades were oscillated about the mid-chord, while the third blade is chosen as the reference. ANSYS CFX is used for computing the transient solution.

The effect of the phase difference between the neighboring blades is referred to as interblade phase angle (IBPA). The reduced frequency is defined as , where b is the semi-chord, ω the blade oscillation angular

frequency and U is the inlet velocity. The damping parameter is a function of the phase difference between the blade force component (lift or drag) and the blade oscillation motion.



21

AEDSS-2014-203

3. Results Figure 1 shows the plot of various parameters as a function of IBPA. The phase angle determines whether the conditions lead to instability, while the corresponding amplitude indicated its extent. It can be seen that the amplitudes and phase differences of the forces depend strongly on IBPA while the dependence is minimal for the presently considered range of reduced frequency.

(a) (b)

(c) (d)

Figure 1: Variation of the following parameters with IBPA for different reduced frequencies (a) lift

amplitude, (b) lift phase difference, (c) drag amplitude, and (d) drag phase difference. For the case of the constant blade-to-blade phase angle, instability occurs when the phase difference is between 0° and 180°. In the present case, both lift and drag show a reversal in stability around IBPA = 0°.

4. Conclusion

A moving boundary model has been implemented and tested in ANSYS CFX. Simulations have been

conducted for three values of reduced frequencies and four IBPAs. Regions corresponding to high and low damping for axial (drag) and transverse loading (lift) have been identified using the phase difference information of the periodic loads.

5. References

[1] Carta, F. O. (1982). An experimental investigation of gap-wise periodicity and unsteady aerodynamic response in an oscillating cascade. 1: Experimental and theoretical results (turbine blades). NASA Contractor Rep., CR-3513.

[2] Bölcs, A., & Fransson, T. H. (1986). Aeroelasticity in turbomachines: Comparison of theoretical and experimental cascade results. EPFL.

-150 -100 -50 0 50 100 1500

0.02

0.04

0.06

0.08

0.1

0.12

0.14

IBPA, °

Lif

t A

mp

litu

de

, (N

)

k = 0.255

k = 0.382

k = 0.510

-150 -100 -50 0 50 100 150-180

-90

0

90

180

IBPA, °

Lif

t p

ha

se

dif

fere

nc

e, °

k = 0.255

k = 0.382

k = 0.510

-150 -100 -50 0 50 100 1500

0.5

1

1.5

2

2.5

3

3.5x 10

-3

IBPA, °

Dra

g A

mp

litu

de

, (N

)

k = 0.255

k = 0.382

k = 0.510

-150 -100 -50 0 50 100 150-180

-90

0

90

180

IBPA, °

Dra

g p

ha

se

dif

fere

nc

e, °

k = 0.255

k = 0.382

k = 0.510


22

AEDSS-2014-204

Computation of Supersonic Flow Past Backward Facing Step in OpenFOAM

Rahul Kumar Soni, Nitish Arya† and Ashoke De

*Phd Student, IIT Kanpur, Kanpur, 208016. †M.Tech Student, IIT Kanpur, Kanpur, 208016.

Assistant Professor, IIT Kanpur, Kanpur, 208016. Corresponding author‟s email: [email protected]

Corresponding author‟s Ph. No.: +91-512-2597863

Extended Abstract

The understanding of flow field associated with Scramjet engine is crucial in the efficient design of high speed propulsion system. The flow physics is often complex and experimental study with prototypes are quite expensive. This immediately poses the need for robust and customizable solver that could be utilized to investigate detailed flow structure. In the present work, systematic validation of density based solver (rhoCentralFoam) available in OpenFOAM, for supersonic flow, is carried out. The test case investigated here is backward-facing step at Mach 2, reasonable agreement was observed in numerical and experimental results.

1. Introduction

Scramjet engines are found efficient air-breathing propulsion system in high speed flow regime. However, due to intrinsic difficulties associated with the combustion mechanism, there exists need for development in the area of fuel mixing and flame holding. At higher supersonic speed, due to very small residence time (O~ms), achievement of efficient fuel mixing and flame holding is of paramount importance. In recent years backward-facing step being conventional geometric configuration to establish subsonic recirculation zone, has been studied widely by many researchers [1-3]. They also found out that low recirculation zone established behind the step prolongs the residence time. Huang et al [3] put forward that vortices generated at the step corner enhance the fuel and air mixing which is reflected in the improved combustion and mixing efficiency.

The objective of the present work is to evaluate the capability of OpenFOAM framework, an open source computational fluid dynamics class library based on C++ [4], for high speed computation. The density based solver, that utilizes Kurganov and Tadmor schemes is chosen and modified as per our needs to simulate the flow physics over backward-facing step at Mach 2. The test case investigated here represents the experimental investigation of McDaniel et al [2].

2. Formulation

Numerical results were obtained by employing density based solver (rhoCentralFoam) available in OpenFOAM. Compressible, unsteady mass averaged, Reynolds-averaged Navier-Stokes equation are solved with cell-centered finite volume scheme [5]. Turbulence is represented through one equation Spalart-Allmaras (SA) and two equation models like, k-ε, RNG k-ε and SST k-ω. A comparative study of these turbulence models is also performed as part of current work. The molecular dynamic viscosity was evaluated by Sutherland's law, turbulent and molecular Prandlt number is 1 and 0.7 respectively. The convection terms are discretized using monotone preserving schemes and diffusion terms are discretized using central difference scheme, temporal discretization is obtained through second order backward scheme. At inlet boundary uniform flow properties, i.e. Mach number (M∞= 3), free stream velocity (V=520 m/s), static pressure (P∞ = 35 KPa) and static temperature (T∞ = 167 K) are specified. Adiabatic, no-slip boundary conditions were enforced at the top and bottom wall along with condition that noraml pressure gradient vanishes at wall

3. Result The numerical results computed through various RANS models are validated against the experimental results. The streamwise velocity, static temperature and pressure profiles at two locations, namely, x/h=1.75 and 6.66,



23

AEDSS-2014-204

are compared with experimental data, presented in Figure 1. It is observed that numerically predicted results match well with experimental data at both the locations for almost all the turbulence models, except in the region of y/h < 1, which is consistent with the observations of Huang et al. [3].

Figure 1 : Comparison of numerical results with experimental observations at various

streamwise locations

1. Conclusion Systematic validation of OpenFOAM for supersonic flow has been successfully attempted. Flow at Mach 2 over backward-facing step was studied and validated. Reasonable accuracy was observed, except for pressure profile in the region y/h < 1, due poor performance of most of the RANS based turbulence models in the separation region. Minor oscillation after the reattachment shock was observed which needs to be addressed. Overall OpenFOAM, appear to be competitive computational fluid dynamics tool for the investigation of supersonic flow physic

2. References

[1] Takahashi S, Yamano G, Wakai K, Tsue M, Kono M (2000) Self-ignition and transition to flame-holding in a rectangular scramjet combustor with a backward step. Proc Combust Inst 28:705–712 [2] McDaniel JC, Fletcher DG, Hartfield RJ (1992) Staged transverse injection into Mach 2 flow behind a rearward-facing-step: a 3D compressible flow test case for hypersonic combustor CFD validation: AIAA Paper 1992-0827 [3] W. Huang, M. Pourkashanian, L. Ma, D. B. Ingham, S. B. Luo, and Z. G. Wang, “Investigation on the flameholding mechanisms in supersonic flows: backward-facing step and cavity flameholder,” Journal of Visualization, vol. 14, no. 1, pp. 63–74, 2011. [4] Weller, H. G., Tabor, G., Jasak, H., and Fureby, C., "A tensorial approach to computational continuum mechanics using object-oriented techniques," Comput. Physics, Vol. 12, No. 6, 1998, pp.620–631. [5] Christopher JG, Henry GW, Luca G, Reese JM. Implementation of semi-discrete, non-staggered central schemes in a collocated, polyhedral, finite volume framework, for high-speed viscous flows. Int. J. Numer. Methods Fluid 2010;63(1):1–21.


24

AEDSS-2014-205

Numerical Investigation of Soot Formation in Turbulent Diffusion Flame

Manedhar Reddy B and Ashoke De

* M. Tech. Student, IIT Kanpur, Kanpur, 208016. Assistant Professor, IIT Kanpur, Kanpur, 208016. Corresponding author‟s email: [email protected]

Corresponding author‟s Ph. No.: +91-512-259-7863

Abstract

The present work is aimed at examining the ability of semi-empirical soot models in predicting soot formation in 'Delft flame III', which is a non-premixed pilot stabilized natural gas flame. The two-step model is a semi empirical model, where the soot formation is modelled by solving the governing transport equations for the soot mass fraction and normalized radical nuclei concentration. The turbulence-chemistry interaction effects on soot formation are studied using a single variable probability density function (PDF) in terms of a normalized temperature. The results shown in this work clearly elucidate the effect of radiation and turbulence-chemistry interaction on soot formation. The soot volume fraction decreases with the introduction of radiation interactions, which is consistence with the theoretical predictions. It has also been observed in the current work that the soot volume fraction is sensitive to the variable used in the PDF to incorporate the turbulence interactions.

1. Introduction

Pollutants are not the unburned hydrocarbon, they are the hydrocarbon produced during combustion and not consumed by the flame [1]. Soot is an impure form of carbon particles resulting from the incomplete combustion of hydrocarbons which indicates poor utilization of fuel. Nucleation is the starting step of soot formation. It is followed by surface growth of polycyclic aromatic hydrocarbon (PAH) species and results in formation of small particles. The growth of these small particles depends on the heterogeneous surface reactions with acetylene, which is one of the primary growth species. Modeling of these reactions is done by H-Abstraction-Carbon-Addition (HACA) mechanism [2]. The collisions between this small particles result in coagulation. Heterogeneous reactions with molecular oxygen and OH radicals result in the oxidation of soot. In addition to the type of fuel, soot formation also depends on the operating and prevailing flow conditions. Therefore, the soot formation can be reduced by controlling operating parameters like residence time, temperature and the turbulence.

2. Formulation

The turbulence-chemistry interactions are modeled using a steady Laminar flamelet method (SLFM) and presumed shape Eulerian PDF Transport method. The radiative transfer equation is simplified based on the optically thick medium approximation into a truncated series expansion in spherical harmonics (P1 approximation).

Soot Modeling

In the one-step model, a single transport equation is solved for the soot mass fraction Eqn (1)

·( ) ·( )soot tsoot soot soot

soot

YY Y R

t

(1)

Where Rsoot is the net rate of soot generation, calculated using the balance of soot formation and combustion. In the Two- Step model, in addition to the above Eqn 2, the normalized radical nuclei concentration is also solved

** * *·( ) ·( )nuc tnuc nuc nuc

nuc

bb b R

t

(2)

In the Moss-Brookes model, the soot mass concentration is calculated instead of Rsoot in Eqn 1 and the instantaneous production rate of soot particles is calculated instead of R*

nuc in Eqn 2.



25

AEDSS-2014-205

Burner Description

The burner consists of an axisymmetric jet centred in an annulus in which a number of premixed flames are used to stabilise the flame. The fuel is natural gas and is assumed to consist of 81% CH4, 4% C2H6 and 15% N2 by volume.

Numerical Details

In the current work, all computations are performed using ANSYS FLUENT 13.0. For all equations, a second order discretization scheme is used. SIMPLE algorithm is used for pressure-velocity coupling.

3. Result Contours of mean temperature indicate that with the inclusion of gray radiation the peak centerline temperature drops by ~100K and with the non-gray radiation the peak centerline temperature dropped by another ~350K and the peak temperature now shifted by ~12D in the upstream direction. Radiative transfer of thermal energy from the flame is substantial increased by soot which has high emissivity. This has lowered the flame temperature. The soot volume fraction has reduced by a factor of ~2 with inclusion of gray radiation and by a factor of ~10 with inclusion of non-gray radiation.

Figure 1: Contours of mean temperature and soot volume fraction with (a) no radiation, (b) gray radiation and (c) non-gray radiation using EPDF for TCI.

4. Conclusion a. The gray radiation was not sufficient to account for the radiation losses and the non-gray model has

significantly lowered the global maximum temperature by ~300K and temperature was found to be ~1900K.

b. The soot volume fraction was observed to be very sensitive with temperature and reduced significantly with the gray radiation. But still over predicted the experimental value by a factor of ~10.

c. The soot nucleation is highly sensitive to the radial maximum temperature and reduced by a factor of 4 with the inclusion of non-gray radiation. The soot surface growth and oxidation also show a similar trend with the inclusion of non-gray radiation.

5. References

[1] Haynes, B. S., and Wagner, H. G., 1981. Soot Formation. Progress in Energy and Combustion Science, 7, pp. 229-273. [2] Frenklach, M., and Wang, H., 1990. Detailed Modeling of Soot Particle Nucleation and Growth. 23rd Symposium (International) on Combustion, The Combustion Institute, pp. 1559–1566. [3] Bart Merci, Bertrand Naud, and Dirk Roekaerts, 2005. Flow and Mixing Fields for Transported Scalar PDF Simulations of a Piloted Jet Diffusion Flame („Delft Flame III‟), Flow, Turbulence and Combustion, 74, pp. 239–272.


26

AEDSS-2014-206

Effect of air to liquid momentum flux ratio on primary jet breakup in cross flow of air at atmospheric pressure

Deepak Kumar and A. Kushari†

*M. Tech Student of Aerospace Department, Kanpur, Uttar Pradesh, 208016. † Associate Professor of Aerospace Department, Kanpur, Uttar Pradesh, 208016.

Corresponding author‟s email:[email protected] Corresponding author‟s Ph. No.: +91-941505712

Extended Abstract

This work represent experimental study of primary breakup of liquid jet in an annular passage in a cross flow of air at a fixed mach number 0.12 at ambient pressure. The experiment is conducted for various velocities of liquid jet from 1.417 m/s to 7.084 m/s (based on orifice diameter = 1mm). The droplet distribution (diameter and velocity) and volume flux is measured by Phase Doppler Particle Analyzer in three mutually perpendicular direction. High speed images of spray are captured at the frame rate of 1000 fps. Observed droplet sizes and velocities distribution show that there are three distinct zones. However, different trend is observed in all zones. Furthermore, the dimensionless number (Weber number and Reynolds number) contribute only during breakup process and becomes independent later.

1. Introduction

Liquid jet atomization in cross flow of air has major applications in propulsion systems such as gas turbine engines, ramjet and scramjet. The combustion efficiency and NOx emissions from these engines are strongly dependent on fuel/air mixing ratio, spatial distribution of liquid drops and drop vaporization rate. This requires detail understanding of jet breakup. The breakup of liquid jet in cross flow of air is caused by the multiphase force interactions like inertial force of air, viscous force of air and liquid and surface tension force of liquid. These processes affect the atomization process in a complicated way. Present work study, experimentally, the effect of jet momentum ratio on jet breakup process.

2. Experimental setup

Experimental set up (see Figure 1) consists of a pressurized cylindrical chamber containing test section leading to air- liquid cross flow. Inlet air velocity was measured using a pitot static probe. Phase Doppler Particle Analyzer (PDPA) was used to measure drop sizes, axial and radial velocities for entire spray. High speed camera was used to capture images at 1000 fps and LED light for illuminating the region of interest. Water flow rate was measured using a turbine flow meter. Centrifugal blower was used to create air flow rate.

Figure 1. Experimental set up



27

AEDSS-2014-206

3. Result The dimensionless number (Weber number and Reynolds number), volume flux and jet velocity at various momentum flux used during experiments are tabulated in Table 1. Figure 2 shows sauter mean diameter variation at various momentum flux along axial direction. Three distinctive zones (Zone I:4-9mm, Zone II: 9-19mm, Zone III: 19-29mm) are recognized where sauter mean diameter decrease with axial direction. Small size drops are present in zone III, where the deviation more for different momentum flux. In zone II (medium size drops) the change is moderate at various momentum flux. However at zone I relatively large size drops are present which are less affected by momentum flux.

Figure 2. variation of Sauter mean diameter with axial distance

q 5.00 7.00 10.00 12.00 15.00 16.00 18.00 20.00 23.00 25.00

Vj (m/s) 3.17 3.75 4.48 4.91 5.49 5.67 6.01 6.34 6.79 7.08

Re 3168.15 3748.60 4480.44 4908.07 5487.39 5667.35 6011.14 6336.30 6794.92 7084.19

We 139.60 195.44 279.20 335.04 418.80 446.72 502.56 558.40 642.15 697.99

Q 149.30 176.65 211.14 231.29 258.59 267.07 283.27 298.59 320.20 333.83

Table 1. various air to liquid momentum flux ratios and corresponds to dimensionless numbers.

4. Conclusion Penetration height increases as q increases. Droplet size is constant along radial direction and reduces in axial direction after jet breakup. Axial velocity increases with axial direction at given radial distance and increases with lower axial distance with radial direction. Radial velocity decreases along axial direction and increases along radial direction. Weber number and Reynolds number of drops contribute only in breakup process.

5. Acknowledgement

Authors would like to thank Pratt and Whitney, USA for sponsoring this research activity.

6. References

[1] Wu PK, Kirkendall KA and Fuller RP, 1997, Breakup processes of liquid jet in subsonic crossflows, Journal of Propulsion and Power, 13(1), 173-182.

[2] Santolaya JL, Aisa LA, Calvo E,Garcia I, Garcia JA,2010, Analysis by droplet size classes of liquid flow structure in a pressure swirl hollow cone spray, Chemical Engineering and Processing, 49(1 ), 125-131.

[3] Elshamy OM, Tambe SB, Cai J, Jeng SM, 2007, PIV and LDV measurements for liquid jets in cross flow, AIAA

2007-1338, 45th AIAA Aerospace Sciences Meeting and Exhibit, Reno, Nevada

200

250

300

350

400

450

500

4 9 14 19 24

D32(µ

m)

Z(mm)

Sauter mean diameter Vs axial distance

q=4,y=2mm

q=6,y=3mm

q=9,y=5mm

q=11,y=5mm

q=13,y=5mm

q=16,y=6mm

q=20,y=7mm

q=23,y=7.35


28

AEDSS-2014-207

Active Control of Hooting in Gas Turbine Engines

Dinesh Kumar.S.J. and A. Kushari†

*Ph.D Scholar of Aerospace Department, Kanpur, Uttar Pradesh, 208016.

† Associate Professor of Aerospace Department, Kanpur, Uttar Pradesh, 208016.

Corresponding author‟s email:[email protected]


Extended Abstract

The present experimental study investigates the combustion instability phenomenon in a Gas turbine

combustor which employs swirling flow of air and an Air-blast atomizer to inject both fuel and Atomizing

air. The Combustor employs four air supplies (Atomizing air, Swirling air, Quenching air and Secondary

air). For this present study swirling air, secondary air and quenching air flow rate is fixed and only the

atomizing air flow rate is varied, Quenching air was reduced in phases and the reduction of quenching

air was compensated by primary air. The data indicated that the predominant frequency of oscillations is

around 300Hz which is twice that of designed value (150Hz). Combustion instability is predominant with

45 degree swirler and the magnitudes of oscillations in the combustor are matched with that of acoustic

measurements.

1. Introduction

Combustion instability remains one of the critical issues in developments of any combustion systems. The

phenomenon is observed in all types of propulsion systems. Stringent emission levels have restricted the

operation of Diffusion controlled Gas turbine combustors. So to control the NOx formation, development of Lean

premixed combustor has started. But these combustors are prone to get combustion instability phenomenon as

they operate in very lean conditions. Hence, lean combustion process generally exhibits disturbances in flow

parameters such as pressure, temperature, velocity, and species concentration. If these disturbances interact

with the heat release rates then the disturbances grow with time and they form limit-cycle oscillations having

large amplitude. This growth of disturbances due to their interaction with the combustion process (heat release

rates) is referred as Combustion Instability.

2. Experimental setup

Experimental set up (see Figure 1) Air blast Atomizer shown in figure 1(a), shows that air enters the atomizer tangentially and moves into a manifold (which consists of 1mm hole of eight numbers) and from that manifold it enters into the main manifold of 11 mm inside diameter. Fuel enters axially and leaves the fuel pipe radially. Air which comes out of the manifold pushes the fuel flow out from the Air blast Atomizer. A 400 notch angle is provided at the exit so that it helps in better mixing of the air and fuel with the swirling air.



29

AEDSS-2014-207

Figure 1. Schematic of Experimental set up (Left Air Blast Atomizer)

3. Result

Quenching air was found out to be one of the dominating sources for hooting in gas turbine, Hence by

decreasing the quenching air, It was found that the SPL values were decreased upto 25db even at lean

combustion process. By pulsing down the atomizing air at 15hz with an solenoid valve it was found out that

the instability decreases due to change in the droplet diameter with the apparent changes in the coherent

structures across the dump combustor.

Pressure and Heat Release Fluctuations FFT‟s of Corresponding Pressure and HR

Fig.1 Details of Pressure and Heat Release at FAR 0.0015


30

AEDSS-2014-207 4. Conclusion

Most of the tests showed frequency of oscillations as 360Hz, which was supported by the acoustic

measurements. With 45 swirler a sound pressure level of 145db was achieved and was reduced by 26db by

decreasing the quenching air.

5. Acknowledgement

The authors would like to acknowledge the sponsors Pratt & Whitney (P&W) Cannada, under the project of Active and Passive Control of Hooting in Gas Turbine Engines.

6. References 1. Sreenivasan, K. R., and Raghu, S., " The control of combustion instability: A perspective ", Vol. 79,

September 2000.

2. Jakob J.Keller., “Thermo-acoustic oscillations in Combustion chambers of Gas Turbines”. AIAA , Vol.33,

No.12 , December 1995.


31

AEDSS-2014-301

ADDITION OF A LEAD LAG DAMPER MODEL AND CHANGE OF ROOT BOUNDRY

CONDITIONS IN COMPREHENSIVE AEROELASTIC CODE AND ITS EFFECTS ON

STRUCTURAL DYNAMICS AND BLADE LOADS

Parwez Alam, Dr. C Venkatesan†

*Graduate student, [email protected]

† Professor and former Head, [email protected] Department of Aerospace Engineering

Indian Institute of Technology, Kanpur, India, 208016

Abstract

An Analytical linear model of lead lag damper is developed by accounting the flap, lag and torsion kinematic coupling of damper attachment point of blade. The lead lag damper is mounted between the rotor blade and rotor hub. This damper model is added to the comprehensive aeroelastic code for finding the natural frequency and mode shape of hingeless rotor blade which is very much useful to calculate the loads on the blades. In structural modeling rotor blades are treated as long slender beams undergoing deformation in flap, lag, torsion and axial. The governing equation of motion of rotor blade is derived using Hamilton’s principal. The change of root boundary condition (blade attachment to the hub ) in aeroelastic code is also analyze. With the change in root boundary condition the loads on the blade correlates very well with industry results.

1. Introduction

Helicopter operates in a very complex dynamic and aerodynamic environment. The complex loading environment of the helicopter is due to the continuous interaction of aerodynamic, structural, centrifugal and inertia forces acting on it. Helicopters with articulated and soft in plane hingeless rotor are known to be susceptible to aeromechanical instabilities such as ground and air resonance [2]. These instabilities arise due to the coupling of the poorly damped rotor cyclic modes with fuselage modes. The ground and air resonance instabilities cause uncomfortable and dangerous conditions for the helicopter and its occupants. In the case of ground resonance, the instability can result in the complete destruction of the aircraft [3]. To eliminate the danger of these instabilities, root end lead lag dampers have been used.

2. Formulation

The coupled equation of motion for rotor blade is derived by using Hamilton‟s variational principle. Damper kinematic equation is derived using the variation of strain energy stored in the damper.

The generalized Hamilton principal is expressed as equations

Where δU is the virtual variation of strain energy, δT is the virtual variation of kinetic energy and δWe is the virtual work done by the external forces[1] .

2

1

( ) 0

t

e

t

U T W dt


32

-8000

-6000

-4000

-2000

0

2000

4000

6000

-2 0 2 4 6 8 10 12 14

Flap

mo

me

nt@

0.4

35

m(N

m)

[email protected](Deg.)

analysis

whirl tower

-5000

-4000

-3000

-2000

-1000

0

1000

2000

-2 0 2 4 6 8 10 12 14

Lag

mo

me

nt@

0.8

m(N

m)


analysis

whirl tower

-20000

0

20000

40000

60000

80000

100000

-2 0 2 4 6 8 10 12 14 16 18

Thru

st


analysiswhirl tower

0

500

1000

1500

2000

-2 0 2 4 6 8 10 12 14 16 18

Po

we

r


analysis

whirl tower

AEDSS-2014-301

3. Result The performance characteristic of a rotor is validated with whirl tower test data from industry. Figure 1 shows the thrust and power curves. Figure 2 shows the variation of root bending moment in lag and flap direction with respect to collective pitch angle of the rotor blade. It can be seen that analysis result match with whirl tower test data fairly well.

4. Conclusion

The loads on the blade from analysis were found to correlate well with the whirl tower test data.

5. References

[1] Rohin Kumar M , “ Comprehensive Aeroelastic and flight Dynamic Formulation for the Prediction of Loads and Control Response of a Helicopter in General Maneuvering Flight ,” Ph.D Thesis, Dept. of Aerospace Engineering. I.I.T Kanpur, Feburary. 2014 [2] Conor Marr, “Conceptualization, Modeling and Characterization of a CF Driven Multi State Lead Lag Bypass

Damper,” Ph.D. Thesis, Dept. of Aerospace Engineering, Pennsylvania State University, May 2012. [3] Byers, L.K. , “ Helicopter Rotor Lag Damping Augmentation based on radial absorber & Coriolis Coupling,” Ph.D. Thesis, Dept. of Aerospace Engineering, Pennsylvania State University, August 2006.

Figure-1: Thrust and power curves for whirl tower case

Figure 2: Bending moments variations with collective angles


33

AEDSS-2014-302

Development and Structural Dynamic Analysis of Bio-inspired MAV Flapping Wings

David Kumar, Vemuri Shyam, Tigmanshu Goyal, P M Mohite, Sudhir Kamle

Indian Institute of Technology Kanpur, Uttar Pradesh, 208016. Corresponding author‟s email: [email protected]


Extended Abstract

The aim of this work is to develop and apply polypropylene (PP), carbon nanotubes (CNTs) composites, for making micro air vehicle (MAV) flapping wings. MAVs (small unmanned aircrafts) can be used for surveillance, reconnaissance, biochemical sensing, targeting, tracking, etc. [1]. While they can be fixed wing, or rotary wing, flapping wing MAVs have the advantage of being able to fly at low speeds and have high maneuverability. They generate both lift and thrust using their wings only. The development of a flapping wing MAV presents many technical challenges. Some challenges include materials and structural aspects, low Reynolds number aerodynamics, weight and volume constraints, high propulsive power to weight ratio requirements and ability to adapt to all environments. MAVs require to do hovering for most of the missions. Hummingbirds (Giant Hummingbirds), having excellent flight (mainly hovering and backward motion) and structural characteristics (make reverse camber to generate lift in both strokes) [2], were chosen as the bio-inspiration for wing development. The wings are required to be light, strong and fatigue resistant, to perform flapping motion during flight. Therefore, wing-material becomes a crucial component. An optimization analysis, on the basis of density and fundamental frequency obtained through Ansys, was done for selecting wing material. Polypropylene, a thermoplastic polymer, was observed to have desired properties such as light weight, flexibility, good strength, fatigue resistance, good heat and chemical resistance etc. Mixing CNT with PP further increased the strength significantly, making it more suitable for MAV application. Xylene was used as the solvent for making PP-CNTs composite. The films were developed using two flat plates and putting them in UTM thermal chamber, at 230-240 ºC, under compression mode. The fabricated films (0.1-0.2 mm thin) were characterized mechanically (using UTM) and thermally (using TGA and DSC). 20% increment in strength of PP was observed at 0.1 wt% of CNT. At CNT concentration higher than 0.1 %, the strength of PP decreased to significant level. There was no improvement or degradation observed in thermal properties. The results repeatability is one of the key factors of manufacturing techniques. One must be able to manufacture the same object with same properties repeatedly. Here, we need at least two wings for MAV application and with same characteristics to avoid uncontrolled asymmetric flapping. A mold was designed to cast the desired wings. The mold for the polypropylene wing was made out of brass from the 3-Axis CNC machine. For fabrication of wings the mold, with the material within it, was put in the thermal chamber of UTM in compression mode. The temperature of chamber was set up to 230-240 ºC. The molds are compressed till they completely fit as designed. The wings were characterized by their structural dynamic characteristics. Wing modal analysis was done to obtain natural frequency and mode shapes. They were observed through stroboscope and recorded using high speed camera. The analysis was aimed to get the fundamental mode in the flapping range (8-15 Hz) of hummingbirds [2]. The testing was also done inside vacuum chamber to observe the effect of air on the natural frequency and modes. The Ansys results were compared with the experiments in vacuum for validation of results. Keywords: Hummingbird wings, MAVs, PP-CNT composites, Structural dynamic analysis

References: [1] Svanberg CE., 2008, “Biometric Micro Air Vehicle Testing Development and Small Scale Flapping Wing Analysis”, MS Thesis, AFIT/GAE/ENY/08-M27, AFIT (AU), Wright-Patterson AFB OH. [2] Raney DL and Slominski EC, 2003, “Mechanization and Control Concepts for Biologically Inspired Micro Aerial Vehicles”, AIAA 2003-5345.



34

AEDSS-2014-303

A-Posteriori Error Estimation for Non-Linear Problems

P. M. Mohite, C. S. Upadhyay and Y. D. Jaiswal

*Indian Institute of Technology, Kanpur, U.P., 208016

Corresponding author‟s email: [email protected]


Extended Abstract

A-posteriori error estimate are so called because they are estimated after the computation of FEM solution. These estimates are important as it gives an estimate of the error associated with the FEM solution and thus shows us the reliability of FEM solution. Although there are numerous literatures on FEM but still the studies on error estimation are immature and are under development. Many commercial codes which are based on FEM are usually complicated and unreliable, further the algorithm employed in solving these problems are often unknown to the user due to proprietary reasons. Hence, it is important to device some mathematical model to estimate the error incurred a-posteriori in the obtained FEM solution and thus, determine the reliability of FEM and improve the accuracy of obtained solution.

1. Introduction We have constructed and validated, as given in Babuska et al [1], several versions of residual-based error estimator for a cantilever bar with a spring (having stiffness coefficients k0 (N/m2) and k1 (N/m4)). The error problem was constructed using the governing differential equation and solved using the known equilibrated residual method (ERM) and various versions of partition of unity (POU) by taking appropriate shape functions. This problem was solved for extreme values of (k/EA) and also with higher order of approximation. This helps us to understand the nature of the error estimates because the same method can be extended to beams and shells and there we expect similar behavior of error estimator. In order to validate the results, we used the ratio of strain energies of the estimated error with the exact error.

2. Result Result obtained using POU, when error problem was solved for the problem corresponding to the exact solution as x2

,

is shown in Figure 1. The ratio of strain energy between exact and estimated error for few extreme cases with constant 'EA' are shown in Table 1, values close to unity signify good accuracy of estimated error.

Exact Sol.

k0/EA (at tip) (m-2)

k1/EA (at tip) (m-4)

Ratio of Strain Energy as obtained

from

ERM POU

x2

1 1 0.999782 1.022623

1 1000 0.860885 1.147323

1000 1 0.385023 1.245680

1000 1000 0.770046 1.234956

3. References

[1] Babuska I, Strouboulis T, Upadhyay CS, Gangaraj SK and Copps K, Validation of a posteriori error estimators by

numerical approach, International Journal for Numerical Methods in Engineering, 37, (1994) 1073-1123.

0 0.2 0.4 0.6 0.8 1-1

-0.5

0

0.5

1

1.5x 10

-3

Err

or

in A

xia

l d

isp

lace

men

t (m

)

Distance from the fixed end of the bar (m)

Exact Error

Estimated Error

Figure 1. Plot between estimated error and exact error with k0/EA=1000 m-2 & k1/EA=1000 m-4 (at tip)

Table 1. Strain energy ratio between exact and estimated error (element length = 0.0625 m)



35

AEDSS-2014-401

Angle of Attack, Pitch Angle and Glide Angle Modeling at Various Thrust Inputs for a

Powered Parachute Aerial Vehicle

Vindhya Devalla, Om Prakash† and Amit Kumar Mondal

*Doctoral Research Fellow, Dept. Aerospace, UPES, Dehradun, Uttarakhand, 248007. † Professor, Head, Dept. Aerospace, UPES Dehradun, Uttarakhand, 248007.

Doctoral Research Fellow, Dept. Electronics, Instrumentation and Control, UPES, Dehradun, Uttarakhand, 248007. Corresponding author‟s email: [email protected], [email protected]


Extended Abstract

The biggest problem oil and gas industry facing today is the safety and security of the pipelines which contributes to

the increase in the maintenance cost of pipeline directly affecting the cost of fuel. An unmanned powered parachute

aerial vehicle was developed for monitoring the oil and gas pipeline. It is envisioned that the vehicle would follow a

planned trajectory to the target, and FLIR based system would be employed for extracting details about pipeline

leakages, thefts,internal corrosion, internal waxing, etc. In this paper, a 9DOF parafoil simulation was created in the

Matlab/Simulink environment to study the parafoil dynamics and assess the feasibility of the system. Thrust , pitch

angle, glide angle and angle of attack was studied and was studied for the developed model and was compared with

lingards’s model.

1. Introduction

A 9 DOF dynamical model of the parafoil was created to assist in the development of the precision delivery project. It is a

critical component of the parafoil simulation and is essential to the creation of the control and guidance systems. The Powered

Parachute (PPC) is an aircraft which derives lift from a ram-air inflated canopy, under which the fuselage is suspended. The

parachute is inflated by the dynamic pressure of the air flowing into the canopy which has a cross section in the shape of an

airfoil. This process helps the vehicle to create lift. This feature differentiates these parafoils from conventional parachutes

which are used to simply create drag. Powered parachutes have been utilized mostly for recreation activities, but some of the

special properties make them a suitable platform for unmanned aerial vehicle (UAV) and remote sensing applications.

Powered parachutes have existed since 1981 [1]. The concept was introduced at the Sun & Fun aviation event by the

ParaPlane Corporation. They represent aircraft that are somewhere between balloons and fixed wing aircraft when control is

considered as shown in Fig 1. The direction of a powered parachute is controlled by the pilot pushing on either a left or right

steering bar that pulls down on a line attached to the trailing edge of the canopy. The increased drag causes the aircraft to

turn.

2. Formulation

The 9 DOF model is given as in equation 1. The parafoil payload model is modeled as two body system consisting of canopy

mass and payload mass suspended below the canopy using the suspension lines. The steering configuration used is known

as a “fly-bar.” In this design, the Parafoil is connected to the ends of the flybar. This bar can be pulled either side of the

aircraft, changing the direction of the lift and making the aircraft turn. This type of parafoil, payload model uses 9 DOF model.

The separation between the Aerodynamic center of the canopy and the payload center of gravity produces a swinging motion.

These two centers are joined by massless links Rp and Rb, At joint C, resulting in 3DOF rotational and 3 DOF translational

motions [3].




36

AEDSS-2014-401

4

3

2

1

00

00

)()(0

0

B

B

B

B

F

VTRII

TRI

TTMMRMM

TTMRM

c

c

p

b

pcpMp

bcbb

bpFpcpFp

bbbcbb

pMpp

A

p

bbb

cpFpcppFcppFp

G

p

A

p

cbbbb

T

b

G

b

A

b

IIMB

IB

VTMVTMRMMFFB

RMFFFB

)(

)(

4

3

2

1

3. Result

The developed PAV model has been validated using the lingard model. Angle of attack (alpha), pitch angle (theta) and glide

angle (gamma) values obtained are compared with the Lingard values under no thrust conditions

Figure1. Parachute Aerial Vehicle under no Thrust condition

The model is validated using the lingard results by checking the values of angle of attack (alpha), pitch angle (theta)

and glide angle (gamma) of the parafoil with respect to time.

Figure2. Validation of Parachute Aerial Vehicle with Lingard model

The developed PAV model has been simulated at various thrust conditions, keeping the thrust as 4.5N, 7.75N and 10N.

It is observed that the gamma value is zero at the thrust input of 7.75N.


37

AEDSS-2014-401

Figure3 . gamma values for various thrust conditions

Keeping the thrust value 7.75N the alpha, theta and gamma values are seen in figure 4

Figure.4. Parachute Aerial Vehicle with 7.75N thrust

Figure.5. PAV landing

From the figure 5, it can be observed that The theta value, while landing the there is glide angle of approximately 8 degrees.

4. Conclusion

The following cases are considered in the above mentioned model validating along with the lingard report.

Lingard Parachute Arial vehicle (Thrust = 0N)

Alpha_p (deg) 3 5

Theta_p (deg) -9 -10

Gamma(deg) -18 -18


38

AEDSS-2014-401

Gamma value with different thrust conditions

S.No. Thrust (N) Gamma(deg)

1 4.5 -8

2 7.75 0

3 9 7

It is observed that at thrust 7.75N the gamma value is 0, i.e. at level flight. Considering thrust value to be 7.75 N, angle of

attack, pitch angle and glide angles are plotted and the following table is obtained.

Parachute Arial vehicle (Thrust = 4.5N)

Parachute Arial vehicle (Thrust = 7.75N)

Parachute Arial vehicle (Thrust = 9N)

Alpha_p (deg) 10 8 7

Theta_p (deg) 0 8 10

Gamma(deg) -8 0 7 It is observed that as the thrust increases the angle of attack decreases and the pitching angle increases. The gliding angle

indicates that the descending or on steady level flight.

5. References

[1] Damian Toohey, “Development of small parafoil Vehicle for Precision Delivery”. Massachusetts Institute of Technology, 2005

[2] O. Prakash and N. Ananthkrishnan, “Modeling and Simulation of 9-DOF Parafoil-Payload System Flight Dynamics,” AIAA Atmospheric Flight Mechanics Conference and Exhibit, Keystone, Colorado, Aug., 2006

[3] Lingard. J. S, “Ram Air parachute Design” 13th AIAA Aerodynamic decelerator system technology conference, Clearwater Beach, May 1995.


39

AEDSS-2014-402

Flight Dynamic Modeling for Mini Helicopter with Stabilizer bar for Trim and Stability Analysis

T.Sakthivel* and Dr.C.Venkatesan †

*Graduate Student, IIT Kanpur,UP-208016. † Professor,Department of Aerospace Engineering IIT Kanpur,UP-208016.

Corresponding author‟s email: [email protected] Corresponding author‟s Ph. No.: +91-7897768700

Extended Abstract

This study describes the development and validation of a flight dynamic modeling for analysis of trim and stability of a mini helicopter including the effect of stabilizer bar. Designing a helicopter meeting the required handling quality aspects ensures low pilot workload, increased mission effectiveness and improved safety of operations in all weather and visibility conditions. Lack of an indigenous tool for helicopter design industry in India is the motivation behind this study. The model consists of development and integration of sub-models for main rotor, tail rotor, stabilizer bar fuselage, horizontal stabilizer and vertical stabilizer. The simulation model has the capability to analyze any flight condition including level, sideslip, climb, descent or turns. The simulation model can be improved by including dynamic stall and dynamic inflow effects.

1. Introduction

Helicopter design is purely an interdisciplinary activity involving dynamics, aerodynamics, flight control and propulsion. Modeling of helicopter dynamics involves main rotor blade dynamics, its coupling with fuselage and the forces associated with the tail rotor and other lifting surfaces. Compared to large helicopters, mini helicopters have certain key features like variable main and trail rotor rpm, rigid rotor blade, stabilizer bar. The stabilizer bar provides the passive control pitch input to main rotor blade. This feedback control pitch input depends on the flapping response of the stabilizer bar, which is influenced by the perturbational motions of the fuselage and also due to cyclic pitch input from the swash plate.

2. Formulation

In this work formulation part includes the force and moment calculation from the various lift generating parts in the helicopter like Main rotor, Tail rotor, Fuselage, Stabilizer bar, Horizontal stabilizer and Vertical stabilizer. The main rotor forces contributed by the inertial and aerodynamic loads. Inertial loads are formulated from Newton law of motion. Aerodynamic forces are formulated from Blade element theory. Sectional loads moments are calculated and integrated over the length of the blade to get net force and moment. Most simplified models are taken into account for force and moment calculation. For the main rotor no lead lag and twisting of the blade taken into account. For tail rotor only thrust is calculated and other forces are neglected. Stabilizer bar flap dynamics is considered and the forces and moment are not taken into account. Fuselage drag is calculated from the equivalent flat plate area. Empennage surface are generating only forces normal to their surfaces. For the trim analysis 6 equilibrium equations are solved for the control variables. Using small perturbation theory, the stability of the mini helicopter is studied.

3. Result The mathematical developed in the previous section is simulated for the mini helicopter of 10 kg weight and 0.91m of blade radius with the stabilizer bar.The control angles to trim the helicopter for various forward speeds are plotted.


40

AEDSS-2014-402

Figure 1. Trim Control angles for various forward speeds

4. Conclusion Trim and stability analysis are performed for the mini helicopter with stabilizer bar and without stabilizer bar.By comparing both the results we can conclude that stabilizer bar will increase the stability of the helicopter. Without stabilizer bar helicopter is unstable at hover and small forward speeds. After including the stabilizer bar that unstable modes are brought to neutrally stable.

5. References

[1] Venketesan C, 1997, Short course on Helicopter Technology Department of aerospace Engineering, IIT Kanpur [2] Gagan Deep Singh, 2012,Helicopter flight dynamics simulation for analysis of trim, stability and control response, IIT Kanpur.

0

4

8

12

16

0.00 0.08 0.16 0.24

TR Collective (deg)

0

6

12

18

0.00 0.08 0.16 0.24

MR Collective (deg)

-4

-2

0

0.00 0.08 0.16 0.24

MR Longitudinal Cyclic (deg)

-8

-4

0

4

8

0.00 0.08 0.16 0.24

MR Lateral Cyclic (deg)

-20

-15

-10

-5

0

0.00 0.08 0.16 0.24

H/c Pitch Attitude (deg)

-12

-8

-4

0

0.00 0.08 0.16 0.24

H/c Roll Attitude (deg)


41

AEDSS-2014-403

Influence of Main Rotor downwash on the Horizontal Tail and its effect on Trim of the Helicopter

Aravindhan D and Dr.C. Venkatesan†

*Graduate student, IIT Kanpur, UP, 208016. † Professor, Dept. of Aerospace Engg. IIT Kanpur, 208016.

Corresponding author‟s email: [email protected] Corresponding author‟s Ph. No.: +91-9886634377

Extended Abstract

1. Introduction

The helicopter loads and vibration prediction analysis is simulated using the Comprehensive Aeroelastic Response (CAR) code. For the purpose of calculating downwash effect on the horizontal tail, Peter-He Dynamic wake model is used. Addition of main rotor induced velocities to the horizontal tail with its downwash effective factor and its percentage impingement is considered for the calculation of loads arising out of the empennage for the improvement of helicopter trim conditions.

2. Formulation

The load generation by the horizontal tail is calculated by the following equations. The wake skew angle ( ) is calculated to find the percentage impingement of the rotor wake on the tail plane. The horizontal tail setting angle

( ) is considered for calculating the effective angle of attack ( ) due to the induced velocity component

( .

3. Result The effect of downwash on the empennage of the helicopter was studied for various speeds of the aircraft. The downwash effective factor ( ) is varied from 1 to 2. The increase in downwash effective factor reduces the collective angle (θ0) and longitudinal cyclic control angle (θ1s) whereas it increases

the Roll attitude (φ), pitch attitude (θ) and Lateral cyclic angles (θ1c). The tail rotor collective angle (θTR)

does not vary much with the downwash effective factor. Figure 1 depicts the variation of control angles with downwash effective factor for advance ratio of 0.2. Table 1 shows the variation of Trim angles with respect to downwash effective factor for advance ratio of 0.2 and 0.25.



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AEDSS-2014-403

Figure 1. Downwash effective factor on control angle variation

μ θ1c θ1s θTR

Flt downwash

Flt downwash

Flt downwash

with With out with With out with With out

0.20 2 1.748 2.148 -4.04 -1.662 -2.750 3.9 4.075 3.973

0.25 1.4 1.583 1.863 -5.07 -2.902 -3.757 3.9 4.477 4.355

Table 1.Trim control angles comparison for advance ratio (μ) of 0.2 and 0.25

4. Conclusion The control angles for the trim of helicopter are improved by considering the main rotor downwash on the horizontal tail as is evident from the above results.

5. Acknowledgement

I would like to express my gratitude and regards to my guide Prof. Dr. C. Venkatesan for his insightful guidance throughout this course of work. I would also like to thank Mr. DSD Prasada Rao, DGM, RWR&DC for his support and guidance. This program is made possible by generous support from Hindustan Aeronautics Limited.

6. References

[1] David A. Peters, David Doug Boyd, Cheng Jian He, 1987, Finite-State Induced-Flow Model for Rotors in Hover and Forward Flight, Proceedings of 43rd Annual Forum of AHS, St. Louis, Mo., May 18 – May 20. [2] Rohin Kumar M, 2014, Comprehensive Aeroelastic and Flight dynamic formulation for the prediction of loads and control response of a helicopter in general Maneuvering flight, Doctoral Thesis, Indian Institute of Technology Kanpur, February 2014. [3] Robert T.N. Chen, A Survey of Nonuniform Inflow Models for Rotorcraft Flight Dynamics and Control Applications, NASA Technical Memorandums, Ames Research Center, Moffett Field, California, 64(1) – 64(62), November 1989.

[4] Gareth D. Padfield, Helicopter Flight Dynamics, 2007, AIAA Education Series, 2nd Edition, Blackwell Publishing,

North America.

-5

0

5

10

1 1.2 1.4 1.6 1.8 2

con

tro

l an

gle

s

downwash effective factor

θ φ θ0 θ1c θ1s θTR

μ θ φ θ0

Flt downwash

Flt downwash

Flt downwash

with With out with With out with With out

0.20 1.96 2.058 2.653 -2.24 -0.992 -0.759 7.85 8.187 9.287

0.25 -0.60 0.809 1.288 -2.04 -1.574 -1.374 8.38 8.755 10.552


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AEDSS-2014-404

Estimation of Nonlinear Parameters from Simulated Data of an Aircraft

Dhayalan.R, A.K.Ghosh†

*Doctoral Student, Department of Aerospace Engineering, IIT Kanpur, India.

† Professor, Department of Aerospace Engineering, IIT Kanpur, India.

Corresponding author‟s email: [email protected]


Extended Abstract

The current paper discusses an attempt for estimating Non-linear parameter by an improvement to well-

known Neural Gauss Newton (NGN) method, which makes the method capable of estimating nonlinear

parameters from flight data. The estimation is carried over for a set of simulated data with various

control surface combinations. Then the estimation is carried out for the simulated data with selected

control surface combination, for which noise is added, to test the handling capabilities of the Improved

Neural Gauss Newton (INGN) method.

7. Introduction

The NGN method [3] proposed by Peyada et. al, serves as an excellent tool in recent times for parameter

estimation from flight data. An improvement in NGN algorithm [1] has been tried out, which was proved to be

worth of trying the same base algorithm for estimating nonlinear parameters. The estimation process we used is

mainly based upon the angle of attack pitch rate ( ) and linear accelerations along X and Z axis ( and

). In NGN method, the input variables for training include the above mentioned. The force and moment

coefficients are provided along with the accelerations, angle of attack, pitch rate and pitch angle ( ) as outputs

for training the neural model. The data used for estimation is the trajectory model of the airplane under

consideration i.e. time history of the variables. Since the neural model is developed based on the input and

output vector, the idea of improving NGN is tried by including the first derivative of the important motion

variables, in the longitudinal case and for training. By these two inputs, we can characterize the force and

moment coefficients ( , and ) better than NGN method. The pattern following ability of the neural network

is exploited for this improvement. This improves the robustness of the neural model too, since we give the

derivatives of the most influential variables for training. This improvement has brought better estimation of linear

model with a much lower standard deviation. This is has been validated using the HANSA-3 flight data for

longitudinal case, along with NGN for comparison [1]. The improvement made has been proved very effective for

linear model. With this method, we can estimate nonlinear aerodynamic model.

8. Formulation

The simulation is carried out for a McDonnell Douglas F-4 aircraft, which has nonlinear aerodynamic derivatives

over 3 different angle of attack regions. The nonlinear derivatives include the some of the lateral variables too.

Hence, the data required for estimation process requires the inclusion of trajectory variable angle of side slip. To

generate such data, we have to simulate the data by deflecting both elevator and rudder. By deflecting both

elevator and rudder, the combined effect of angle of attack and angle of sideslip is captured in the data, which is



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AEDSS-2014-404

also carried out for four different combinations of the control surfaces which maximizes the energy of the whole

input[2]. From this the best combination of the control surfaces is chosen and used for estimation of nonlinear

parameters after the introduction of noise. The simulation is carried out for two Angle of attack regimes, i.e.

and .

Result

The estimation results are given below. The plots are presented for the comparison of Linear and Nonlinear parameters over

two angle of attack regions. The Improved NGN method is able to predict the nonlinear parameters from simulated data

without noise with better accuracy. This method is carried out for noise introduction too, which is not presented here.

9. Conclusion


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AEDSS-2014-404

The nonlinear aerodynamic parameters are estimated using the Improved NGN method. The estimation is

carried out for noise-introduced-simulated flight data, which will be presented in the seminar.

10. References

[1] Dhayalan.R and A.K.Ghosh. Improved NGN algorithm for parameter estimation from flight data. In SAROD

2011, 2011.

[2] R. V. Jategaonkar. Flight Vehicle System Identification A Time Domain Methodology. AIAA, Inc., Reston, VA,,

2006.

[3] N. K. Peyada and A. K. Ghosh. Aerodynamic parameter estimation using new filtering technique based on

neural network and gauss-newton method. In ARMS 2008,. ARDE, 2008.


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AEDSS-2014-405

Design, Instrumentation and Data Acquisition of 5 Degree of Freedom (5-DOF) Dynamic Test Rig

Subrahmanyam S1, a, Dhayalan R2, b, Rishi Raj Singh3, c, Ghosh A K4, d.

1,2 Ph.D scholar, Flight Dynamics Lab, Department of Aerospace Engineering, IITK, Kanpur, INDIA.

3 Senior Research Associate, Flight Dynamics Lab, Department of Aerospace Engineering, IITK, Kanpur,

INDIA.

4 Professor, Flight Dynamics Lab, Department of Aerospace Engineering, IITK, Kanpur, INDIA.

[email protected], [email protected]

b, [email protected]

c, [email protected]

d

Keywords: Instrumentation, Data Acquisition, 5-DOF Dynamic Test Rig, 3-DOF gimbal, 2-DOF gimbal, Data telemetry, compensator.

Abstract. The present paper discusses about the design of various components namely 3-DOF

gimbal, 2-DOF gimbal and compensator etc., of 5 degree of freedom dynamic test rig. The paper

also figures out the advantages of the 5-DOF rig in estimating the dynamic derivatives of the

flight vehicle of interest. Importance of the instrumentation of various components of 5-DOF rig

as well as flight vehicle has been explained. It also describes the onboard data acquisition system

as well as data telemetry to ground station which is obtained from the sensors during the tests.

Finally at the end of the paper a sample data is presented which obtained during the wind tunnel

testing of the flight vehicle mounted on 5-DOF dynamic test rig.

Introduction

Any flight vehicle is characterized based upon its aerodynamic derivatives, which

includes the stability derivatives both static and dynamic, and also the control derivatives. Study

of these aerodynamic derivatives of flight vehicle is a part of the design process. Many practical

applications like flight control system design, development of simulators, flying quality analysis

and autopilot design requires the estimation of accurate static, dynamic and control derivatives

[1]. By assuming the flight vehicle as a rigid body the six degrees of freedom (6-DOF) equations

of motion of an aircraft are derived and are expressed below as follows [2].

(1)

(2)

If we observe the above dynamic equations of an aircraft, the total external forces and

moments acting on the flight vehicle is the summation of gravity, propulsive and aerodynamic

forces and moments respectively. These aerodynamic forces and moment coefficients are

modeled as a function of static stability, dynamic stability and control derivatives respectively.

The following first order Taylor series expansion explains the force or moment coefficient as a

function of aerodynamic derivatives [3].

(3)

The above mentioned aerodynamic parameters of flight vehicle can be estimated using

different methods – namely, empirical methods, computational fluid dynamics (CFD) analysis,

wind tunnel testing and, finally from real flight data. In modern engineering, aircraft





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AEDSS-2014-405

development involves a combination of all these methods [4]. Empirical methods, which are

based on the experimental database, are used to estimate the static, dynamic and control

derivatives during the preliminary design of an aircraft [5]. Static parameters are required to

estimate the trim as well as to characterize the handling qualities of an aircraft, where as dynamic

stability derivative enables us to understand how the system responds for external disturbances

over a period of time.

The aerodynamic and stability parameters are usually estimated from the data obtained

through static wind tunnel tests. It is generally difficult to estimate dynamic derivatives through

conventional wind tunnel tests, in which the model is kept static and the data is obtained at each

and every position of interest. Dynamic parameters are in general estimated through flight tests,

which are expensive, time consuming and also fatal during dangerous maneuvers. Further, such

tests are performed on the frozen design, and the scope of modification and design improvements

are very difficult. To minimize the expenditure of real flight testing and to have useful

information at a mid-course of design stage, a 5-DOF dynamic test rig has been designed,

fabricated and tested in the National Wind Tunnel test Facility(NWTF), IITK. The proposed 5-

DOF dynamic test rig is capable for aerodynamic characterization of flight vehicle of any

configuration [6]. The model mounted on this 5-DOF rig can perform roll, pitch, yaw, heave and

sway and the aerodynamic parameters of the flight vehicle can be estimated through variety of

maneuvers. The dynamic data of the flight vehicle obtained from the 5-DOF dynamic test rig is

the time history of the variation of state variables. A 5-DOF mathematical model has been

developed and is used for the simulation of the dynamic motion of the flight vehicle along with

test rig [7]. Parameter estimation techniques, which utilize the generated flight data, are used to

estimate the dynamic parameters of the corresponding flight vehicle [7].

Design and Fabrication.

5-DOF dynamic Test Rig. The flight vehicle of interest, whose dynamic derivatives has to be

determined, will be mounted on the 5-DOF dynamic test rig. Unlike conventional wind tunnel

tests, the model mounted on the 5-DOF dynamic test rig will have the following degrees of

freedom namely, pitch, yaw, roll, heave and sway.

(a) (b)

Fig. 1 (a) Schematic diagram of 5-DOF dynamic test rig indicating various components [5].

(b) Photograph representing 5-DOF dynamic test rig.

The above mentioned DOF are enabled by means of the following components namely

3-DOF gimbal - provides the heave, sway and roll to the aircraft model

2-DOF gimbal - enables the flight vehicle to perform pitch and yaw

Rig Arm - connects the flight model mounted on 2-DOF gimbal to 3-DOF gimbal


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AEDSS-2014-405

Compensator - Excites the system such that the inertial effect of rig arm on model of

interest is negligible.

Instrumentation and Data Acquisition.

In order to acquire the reliable dynamic response of the flight vehicle using dynamic test

rig, proper instrumentation of flight vehicle as well as rig arm is required. Various electronic

instruments and their calibration setups, which are used for data acquisition and data telemetry,

have been explained.

Inertial Measurement Unit (IMU) - Two 9-DOF has been used to extract the linear

accelerations (ax, ay, az), angular rates (p, q, r) and Euler angles (ϕ, θ, ψ) of the flight model as

well as rig arm. This device can be easily mounted on rig arm and flight vehicle model, such that

axis system of device and rig arm/flight vehicle matches. This information is generally used as a

feedback in the control system design[7].

Actuators and Sensors - One of the aspects critical to precise measurement of control

surface deflection and actuation is calibration of both the potentiometers and servo motors, which

are used as sensors and actuators respectively. A calibration rig was designed for this purpose on

which the servo motors and potentiometers can be simultaneously or separately calibrated

(a) Calibration Rig for servo&Potentiometer (b) Calibration plot of Servo (c) Calibration plot of Potentiometer

Fig. 2 Calibration setup and plots for servo and potentiometer

Data Acquisition system - The following block diagram explains the data acquisition

module of the proposed 5-DOF dynamic test rig.

Fig. 3 Block diagram representing the data acquisition system of 5-DOF dynamic test rig.

Wind tunnel Testing.

0 50 100 150 200 250 3000

0.5

1

1.5

2

2.5

Angle (in deg)

Puls

e w

idth

(in

ms)

servo1

servo2

servo3

servo4

0 50 100 150 200 2500

0.5

1

1.5

2

2.5

3

3.5

Angle in deg

Voltage in v

olts

Potentiometer1

Potentiometer2

Potentiometer3

Potentiometer4

Aircraft model

IMU

Potentiometer

Servo

Rig Compensator

IMU

Potentiometer

Servo

Onboard Microprocessor

Data

Telemetry

Data

Logging

Ground Station

GUI Receiver

Data

Transmission

Model Mounted on

2DOF GIMBAL


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The 5-DOF dynamic test rig was tested in National Wind-tunnel Testing Facility

(NWTF), at IIT Kanpur. NWTF is a low speed, closed circuit wind tunnel which is capable of

testing at a wind speed up to 80 m/s. It has the facility to interchange the test sections and the

cross section of each test section is 3 m x 2.25 m [8]. Tests were performed at two Reynolds no.

namely 1.69x105 and 1.13x10

5. The following figure. 4 represents a sample data acquired during

testing.

Fig. 4 Sample data of Euler angles and angular rates acquired during the wind tunnel tests of

5-DOF rig along with flight vehicle model.

References

[1] S.D. Carnduff, S.D. Erbsloeh, A.K. Cooke and M.V. Cook, “Characterizing stability and

Control of subscale aircraft from wind-tunnel dynamic motion,” Journal of Aircraft, Vol.46,

No. 1, pp. 137-147, Jan.-Feb. 2009.

[2] Robert. C. Nelson, Flight Stability and Automatic Control, 2nd

edition, published by TATA

McGraw-Hill companies.

[3] Bandu N. Pamadi, Performance, Stability, Dynamics and Control of Airplanes, Published by

AIAA, © 2004, 2nd

Edition.

[4] D. Greenwell and M. Goman, “Wind tunnel simulation of combat aircraft maneuvers,” Proc.

21st Congress of the International Council of Aeronautical Sciences, No. ICAS-98-3.9.2,

Melbourne, Australia, Sep. 1998.

[5] M.S. Rajamurthy, “Generation of comprehensive longitudinal aerodynamic data using

dynamic wind-tunnel simulation,” Journal of Aircraft, Vol. 34, No. 1, pp29-33,1997.

0 200 400 600 800 1000 1200 1400 1600 1800 2000-5

0

5

ax,

m/s

2

0 200 400 600 800 1000 1200 1400 1600 1800 2000-15

-10

-5

0

az,

m/s

2

0 200 400 600 800 1000 1200 1400 1600 1800 2000

-20

0

20

,

deg

0 200 400 600 800 1000 1200 1400 1600 1800 2000

-20

0

20

q,

deg/s

0 200 400 600 800 1000 1200 1400 1600 1800 2000-2000

0

2000

ea,

deg

0 200 400 600 800 1000 1200 1400 1600 1800 2000

-20

0

20

c(p

itch),

deg

Data Points

Aircraft Data

Rig Data


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[6] A. Gatto and M.H. Lowenber, “Evaluation of a three-degree-of-freedom test rig for stability

derivative estimation,” Journal of Aircrafts, Vol. 43, No. 6,pp. 1747-1762, Nov.-Dec. 2006.

[7] Naba K. Payeda, Manoj K. Dhiman, A.K.Ghosh, “Aerodynamic Characterization of Scale

Model Aircraft using 5-DOF Dynamic Test Rig,” Proc. 2010 NSBE Aerospace System

Conference, Los Angeles, California, Feb. 2010.

[8]Saderla Subrahmanyam, Sunil Sharma and Ghosh A.K, “Analytical Modeling, Trajectory

Simulation and Control of Guided Projectiles,” Proc. Annual International Conference on

Control Automation and Robotics (CAR), Singapore, Feb. 2011.

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