STRUCTURAL TECHNOLOGIES DIVISION

SUMMARY _________________________________________________________________________ The Division was involved in a number of activities including research and technology development, national projects, international projects and inter-disciplinary engineering. In essence, the period saw a holistic development of various activities and interactive working among various groups. In the Dynamics and Adaptive Structures Group (DASG), Ground Vibration Tests (GVT) were performed for various configurations on Gipps Aero GA10 Airvan in Australia to estimate the experimental modal parameters and the flutter behaviour of the aircraft. The buffet qualification for GSLV MkII is the requirement of VSSC (ISRO), in order to ensure that the vehicle is aeroelastically stable in the transonic and in supersonic (at maximum dynamic pressure condition-Qmax) regimes, wind tunnel tests were performed. The active engine mount technology was demonstrated using piezoelectrically actuated active struts. The open and closed loop tests were carried out at the engine test bed, under propeller loadings. The TRL level of 6 is reached for this futuristic technology. The group has developed technologies in the area of adaptive structures for active wing to address vibration control, flutter control & flight flutter testing. Similarly formulation is developed for gust alleviation and ground flutter simulation is demonstrated through reduced order model. Design, analysis and fabrication procedures are addressed for the development of multifunctional composites. Actuator/Sensor de-bonding models are developed in FEM and experimentally validated to design damage tolerant smart structures with AVC/SHM applications. In the Impact and Structural Crashworthiness Group (ISCG), the 8 pound gun was made operational. Design development and fabrication of a forward velocity sled was initiated and is expected to be made operational by June 2013. Work on calibration of gelatin versus real birds for bird strikes was initiated. Apart from this, a crash dummy was utilized to study vibrations in autorickshaws and a low cost seat belt and cushion was devised. In the Fatigue and Structural Integrity Group (FSIG), a major test of the Static Strength tests on Active antenna Array Unit and Life extension of landing gears of a fighter aircraft were completed successfully for customers. Damage tolerance evaluation of a wing root fitting box continued. In addition, progress has been made in the basic understanding of fatigue and fracture mechanics arena, particularly in the area of nano-composites. Several new projects in the area of evaluation of aircraft materials and structures from both government and private industries were initiated. Eleventh five year plan projects were completed and proposals for twelfth five year plan were submitted. Further, Inter-divisional activities continued in the field of mechanical characterization of composite materials, nano-composites, composite repair technology etc. In the Computational Mechanics and Simulation Group (CMSG), it carried out research in computational mechanics, simulation, design and supported work in aircraft programs. Importantly, it has successfully carried out flutter prediction of a transport aircraft and transonic flutter analysis of a combat aircraft wing. The group contributed in flutter prediction for GA10 aircraft and whirl flutter prediction of engine propeller system of the same aircraft. The group carried out research in fluid structure interaction, simulation of riveted lap joints, FE and experimental damage tolerance studies on wing skin panels and MDO studies. Support was provided for LCA, SARAS, NCA and AMCA programs. Reliability studies were conducted by the group on various aircraft systems. The Mechanical Systems Design Group (MSDG), which had its scope expanded commenced work in Design and Development of Mechanical Systems for Aerospace applications such as Landing Gears, Airframe mechanisms, Aircraft Systems etc. The group

also was involved in Advanced / Complex Aerospace model design, Radome Structural Design & Analysis, High Pressure Systems Design & Development, Net shaped technology development for Aerospace application etc. The group recently delivered 1:7.645 scale LCA Mark-2 High speed Air Intake Air Force model to ADA, Bangalore. The model is of complex nature involving more than 230 pressure ports at various regions on external as well as internal duct geometry of the model along the air intake path. A 4.2 diameter GFRP Sandwich Radome for X-band Polarimetric radar was delivered to ISRO-ISTRAC, Bangalore. A water compatible - ultra high pressurizing system of 6900 bar capacity was designed and developed indigenously for CFTRI. Detailed design and analysis of 4000 bar and 6900 bar high pressure vessels was carried out as per BPVC ASME (Sec VIII, Div 3).

1. RESEARCH AND TECHNOLOGY ACTIVITIES

1A. Dynamics and Adaptive Structures Group (DASG)

1A.1 Aeroelastic testing of GSLV MKII The aeroelastic buffet qualification for GSLV MK II is the requirement of VSSC (ISRO), in order to ensure the vehicle safety in the transonic regime. A 1:42 scale model of F06 configuration is designed and ground tested (Figs 1 & 2). Shortly the model will be tested in 1.2 m Trisonic NAL wind tunnel for the required flight conditions to estimate the buffet load on the model.

1A.2 EXPERIMENTAL Aeroelastic Studies on RLV-TD Aeroelastic testing of RLV-TD has been taken up to ascertain the buffeting in transonic and maximum dynamic pressure conditions for the ascend phase (RLV-TD). Also the study of flutter characteristics of the vehicle for the descend phase (TDV) is planned. Models of RLV-TD (1:15 scale) have been designed and analysed, considering the transonic flight condition and maximum dynamic pressure condition in ascend phase. After freezing the design, the transonic model is fabricated and ground tested to establish the model dynamic characteristics in a free-free condition. Further, the model is tested in 1.2m NAL Trisonic wind tunnel and the buffet load in terms of dynamic bending moment distribution is obtained (Fig.3 & 4).

Fig.2: FEM and GVT mode shape comparison (1st bending mode)

Fig.1: GVT of GSLV MK II - F06 model

Fig. 4: BM distribution along the length of the vehicle for different angles of attack –

RLV-TD Transonic model

Fig.3: RLV-TD transonic model in tunnel (Ascent phase – 1:15 scale)

For the descent phase configuration the proposed design of the model (1:8 scale) considers both composite and metallic components. Initially the component level stiffness, mass distributions are verified by analyzing each component separately. Further, integrated model dynamics is obtained through free vibration analysis. A sting is designed to support this model to match the wind tunnel mounting arrangements. Finally, the analysis of complete model on the sting support system (Fig.5) is carried out. Attempt has been made for the critical five modes to be simulated closely to match the frequency requirements (Fig.6a&b). The strength of the model is also examined using the steady loads, supplied by VSSC at component level to ensure model and tunnel safety. The model is currently being fabricated.

1A.3 Use of Multifunctional Materials in MAV’s for Improved Aeroelastic

performances The objective is to develop multifunctional wing for MAV to achieve required trim conditions through adaptive trailing edge. In this project, we have successfully employed Macro Fiber Composite (MFC) actuators on NAL developed BK MAV to realize the hingeless control surface concept through a trailing edge morphing. Numerical analysis, followed by experimental validation have confirmed that the trailing edge morphing technique is useful for achieving the aerodynamic trim conditions for takeoff, level flight etc. The morphed surface can be deployed as elevators (symmetric) and for ailerons (anti-symmetric) in flight for different flight manoeuvres (Figure 7).

Fig.7: Multifunctional structural analysis and testing for trimming (upward and downward continuous camber change)

1A.4 Adaptive Aeroelastic Structures

A 3kg UAV is being designed with suitable wing and tilt rotor concept (Fig.8), which will have vertical takeoff as well as landing capability on the runway. Trailing edge morphing, adaptive winglet and adaptive tail structures are planned. In the first phase, a conventional vehicle will be fabricated, wind tunnel tested and flight demonstrated and subsequently active aeroelastic structures will be integrated to study morphing wing and adaptive tail concept.

Fig.8: Conceptual design of tilt rotor and structural wing design

A generalized unsteady air load approximation procedure using matrix polynomial approach has been developed. This method addresses a better fit for higher order modes (Fig.9) than other techniques for aeroservoelastic problems in state space domain

Fig.9: Unsteady air load approximation using Matrix polynomials approximation

for 4th & 5th Modes

1B. Impact and Structural Crashworthiness Group (ISCG)

1B1. Bird strike Studies: Correlation of test and analysis using synthetic and real birds

In the present study characterization of synthetic bird with that of the real bird and the bird material models available in the explicit Finite Element (FE)-Codes used in the simulation are studied. Attempt is made to approximately replicate the behaviour of real bird to that of the synthetic bird by following various processes of synthetic bird preparation procedures for impact testing and to carry out bird simulation using the material models like Smooth Particle Hydrodynamics (SPH). Test analysis correlation studies were performed by carrying out a series of bird strike test using both synthetic bird and real bird by impacting against rigid targets and on a 1.2mm thick aluminium flat panel at the same velocity and also by validating the simulation results Bird impact testing is done using synthetic bird and with the real bird at an impact velocity of 95m/s at the centre of the flat panel. The impact event is captured using the high speed camera at 5000fps. The impact velocity is measured both using the laser sensor output in the control panel and from the High Speed Video (HSV) using the image processing software tool. The sequence of the impact event both from testing and through simulation results is shown in Fig.10.

Time (ms)

Real Bird Synthetic Bird FE Simulation

0.0

1.0

2.0

3.0

4.0

5.0

6.0

7.0

Fig. 10. Image sequence of bird strike event for real/synthetic bird and simulation bird model at every milli seacond

The impact tested panel both from the real bird and the synthetic bird testing is removed from the test fixture. The permanent deformation contour of the flat panel measured at every 10mm interval throughout the plane is captured using the high precision coordinate measuring machine (CMM). These coordinates are plotted using the pre-processing tool. It was found there is a good correlation of the complete deformation contour and the maximum permanent plastic deformation between the actual, synthetic bird and the FE simulation as can be seen in Fig.11.

(a) Deformed shape of the impact tested panels Deformed shape from

simulation

(b) Real Bird Deformation contour obtained from CMM machine(Max plastic

Deformation of 53.92mm is observed)

(c) Synthetic bird Deformation contour obtained from CMM machine (Max plastic Permanent Deformation of 55.87mm is observed)

(d) SPH Bird Deformation - FE Simulation (Max plastic Permanent Deformation of

58.0mm is observed after 20ms) Fig.11. Deformed shape of the aluminium panel from test and simulation

1B2. Bird strike Studies: Bird strike studies on composite/metal panels

Bird strike tests using synthetic bird and real chicken were conducted on Aluminium 2024-T3 panels of thickness 1.2mm and equivalent composite panels. The weight of the bird was 4 pounds and the velocities were in the range of 80-120m/s. High speed images were taken of the event and the velocities of the bird was found out using image processing software. It was observed that penetration velocity range for 1.2mm thick Aluminium 2024-T3 panel is 117 to 120m/s. The composite panels penetrated at even much lower velocities. Synthetic bird and real chicken gave the same velocity range for the penetration.

1B3. Quasi-static and Low velocity impact studies on gelatin/PMB composite material

Characterization of the Gelatin/Phenolic Microballoon (PMB) composite at quasi static and low velocity impact were carried out with higher water contents (1g:1w, 1g:1.5w,1g:2w). The physical properties like density, viscoelastic behavior and energy absorption of Organic Gelatin/ (PMB) composite was studied with respect to content of Phenolic Microballoon and water content. The physical properties like stiffness and viscous behavior of the composite were studied through stress strain tests at different strain rates (1mm/min 10mm/min 100mm/min and 900mm/min). Low velocity impact tests were conducted to check the energy absorption of gelatin. The low velocity impact tests show that gelatin is a good energy absorber and its energy absorption capability is influenced by porosity (PMB content).

1B4. Installation and commissioning of the Forward Velocity Sled test facility

A bungee operated forward velocity deceleration sled which can generate deceleration pulses as required by FAR 23.562 / 25.562 for aircraft seat qualification testing and for studies on occupant safety for both aerospace and automotive sector is built successfully. The installation and commissioning of the sled is completed and tested for FAR pulses. The sled is first of its kind in the country. The sled is designed for a maximum speed of 60 Kmph and with impact mass of 2200Kg (1200 kg sled + 1000 kg payload). The acceptance tests were carried out to generate the FAR 25.562 crash pulse. Fig.12 shows the image of sled test facility commissioned at Impact and Crashworthiness Research Facility of STTD.

The preliminary trials are being carried out using sled test facility for the evaluation of crash pulses by mounting automotive seat and a human-sac dummy. The tests are also being carried out by adding a simple retro-fit Velcro based seat belt for the human-sac dummy. The purpose of the tests is to see the behaviour of the dummy during the crash event and evaluation of the simple retro-fit Velcro based seat belt. Fig.13 shows the human-sac dummy on the sled with and without seat belt.

Fig.12.Sled test facility

Fig.13 Sac dummy with and without seat belt

1C. Fatigue and Structural Integrity Group (FSIG)

1C1. Carbon fiber composite IM7/ Bismaleimide:

As a part of design data generation and derivation of design allowable, IM7/ Bismaleimide CFC was tested in RT and high temperature conditions. Over 10 different types of

mechanical tests were conducted in RT and 1800C. All the tests were performed following their respective ASTM standards. Over 150 numbers of tests are planned. Testing and evaluation of mechanical properties are underway

Fatigue behaviour of indigenous aluminum alloys: As a part of certification testing, indigenous aluminum alloys developed by DRDO are being tested to determine their tensile and fatigue properties. Aluminum alloy sheet specimens of 2000, 5000 and 6000 series were prepared and tested following their respective ASTM standards. S-N curves generated were compared with MIL standards. Further work on fatigue testing and evaluation is underway.

Evaluation of composite panels under static and fatigue loads: The objectives of the study are to evaluate the tensile and fatigue behaviour of undamaged CFC panel, panel with central hole and a repaired panel. Also, the panels fabricated with honeycomb core and CFC face sheets having similar conditions such as healthy, damaged and repaired were tested under four point bend loading arrangements to evaluate their flexural behaviour. Photographs showing the tensile testing set-up for CFC panels are shown in Fig.14. All the CFC panels were fabricated by ACD. The required strain gages were bonded and the tests were conducted. The load, strain and deflection measurements were acquired through data acquisition system. Further work on testing and evaluation of these panels is underway.

(a) Undamaged (b) Damaged (central hole) (c) Repaired

Fig. 14. Photographs showing experimental setup for tensile testing of CFC panels

Mechanical testing support services: Mechanical testing and evaluation of metallic and composite materials was provided to various other divisions of NAL as well as to outside organizations such as HAL. Tensile properties of Al alloys for C-CADD, Testing support for failure analysis group of Materials Science Division, CAI and open hole compression tests for thermoplastic composites for ACD, Fracture toughness testing and evaluation of Al alloys for HAL etc. were carried out. Research and Development (R&D) Activities: The research and developmental work continued mainly on fatigue behaviour of polymer composites. The experimental evaluation of off-axis fatigue behaviour, prediction of fatigue life under service loads, Mode I delaminations behaviour under spectrum loads, FE analysis of composite panels and comparison with experimental work etc. were carried out. Off-axis fatigue behavior of CFC IMA/ M21:

The fatigue life of a QI and cross-ply CFC under an arbitrary loading angle was estimated from the knowledge of fatigue properties of unidirectional (UD) laminate. IMA/M21 carbon fiber composite UD laminates were fabricated following standard autoclave process. Fatigue test specimens were cut from this laminate so that the fiber to loading axis angle varied from 00 to 900. Static tensile tests were conducted to determine the tensile properties at various off-axis loads. Fatigue tests were conducted to determine S-N curves for the material at various off-axis loading angles (Fig. 15). All the tests were conducted at R =0.1 with a sinusoidal waveform at 3 Hz. Increasing the off-axis angle was observed to decrease the fatigue life. Analytical equations along with the S-N data generated at various angles was used to predict the fatigue life of a quasi-isotropic laminate and compared with experimental results (Fig.16). Further, fatigue life of a cross-ply laminate loaded at an arbitrary loading angle was determined.

Fig. 15. Off-axis fatigue behaviour of IMA/ M21 CFC

Fig.16. Comparison of predicted and experimental fatigue life of QI laminate of IMA/M21 CFC

Mode I delamination onset under spectrum loads:

Standard double cantilever beam (DCB) test specimens of a unidirectional IMA/M21 carbon fiber composite were fabricated using standard autoclave process. A teflon insert of 30 thickness was used to create delamination at the mid plane. Fracture toughness tests were conducted following ASTM test standard specification to estimate the toughness. Based on the load-displacement data obtained during the fracture toughness test, constant amplitude fatigue tests at various maximum displacements were conducted to determine the number of cycles required for onset of delamination growth. All the tests were conducted in a servo-hydraulic test machine under displacement control mode. A photograph of the test set-up is shown in Fig. 17. The onset was determined by monitoring the compliance of the specimen during fatigue test. As per ASTM test standard specification, the onset of growth was considered to be corresponding to 5% deviation from the initial compliance. Fatigue tests were carried out to determine Gmax–Nonset plot at two different stress ratios, R = 0.0 and 0.5. Increasing the stress ratio was observed to increase the onset life. The Gmax–Nonset data was then fit to empirical equation and used to construct a constant onset life diagram (COLD) (Fig.18) which is useful in prediction of onset behaviour under spectrum loads. Further experimental work is underway to validate predicted onset fatigue life under spectrum loads.

Fig. 17. A photograph of test set-up for mode I delaminations studies

Fig. 18. Constant onset life diagram for mode I delamination of IMA/M21 CFC

Spectrum fatigue life of a nanocomposite:

A thermosetting epoxy polymer was modified by incorporating 10 wt.% of silica nanoparticles which were well dispersed in the polymer. Two different glass-fiber reinforced-plastic (GFRP) composite laminates were prepared to give: (i) a GFRP composite with an unmodified epoxy-matrix (termed GFRP-neat), and (ii) a GFRP composite with a silica-nanoparticle modified epoxy-matrix (termed GFRP-nano). Fatigue tests were undertaken employing a standard wind-turbine spectrum-load sequence, WISPERX, shown in Fig. 19. The fatigue life of the GFRP-nano composite was about four times longer than that of the GFRP-neat composite (Fig. 20). This was reflected in (i) the rate of degradation of the stiffness of the composite and (ii) the development of matrix cracking, both being more severe in the GFRP-neat composite, compared to the GFRP-nano composite. The underlying mechanisms for the observed improvement in the spectrum fatigue life of the GFRP-nano composite are discussed. The spectrum fatigue life was predicted following a standard procedure and a good correlation was observed between the predicted and experimental results.

Fig.19: Standard WISPERX load sequence Fig.20. Fatigue life of composites under

WISPERX load sequence

1D. Computational Mechanics and Simulation Group (CMSG)

Structural Design and Analysis support for AMCA Version 3B-08

The AMCA Version 3B-08 structural design, analysis and size optimization is carried out for all critical symmetric and un-symmetric load cases. For the convenience of manufacturing and system installation the fuselage has been divided into three parts viz. Front Fuselage, Middle Fuselage and Rear Fuselage as shown in Fig.20. Joining of two parts will be with continuous fastening during sub assembly stage. Longerons and stringers will not be spliced at the interface of front, middle and rear fuselage location. Wing attached to the fuselage bulkheads at six locations consists two shear and four bending attachment. The boom structure is integral with fuselage and disconnects from fuselage engine rear attachment bulkhead which is over hang. Boom structure supports Vertical and Horizontal tail components. Weapon bay consist central shear wall with box section for supporting bottom doors. Vertical tail attached at three locations, two bending with one shear attachments. H T shaft pivot located in overhang portion of the boom and two Engine attachment bulkheads are provided in the rear fuselage. Bulkheads and longerons are provided in the fuselage to divide skin panels. All bulkheads are distributed at 330 to 350 mm spacing. Fifteen longerons have been introduced and out of these three are continuous from nose to tip and two are continuous along fuselage side shear walls. All longerons are effective for taking bending loads. Local skin stringers will be introduced to break the panels based on detailed stress analysis. T sections for top and channel section for bottom are used for continuous main longerons. Segment longerons of L sections are used at top and bottom of shear wall and boom. Bulkhead frames are attached to the fuselage skin by riveting bulkhead flanges with skin and Longeron flanges are attached to the skin with rivets. Loaded bulkheads are

100

150

200

250

300

350

1.E-01 1.E+00 1.E+01 1.E+02 1.E+03

Ref

eren

ce S

tres

s ,s

ref

(MP

a)

Number of Blocks, Nb

Expt.-GFRP-neatExpt.-GFRP-nanoPred.-GFRP-neatPred.-GFRP-nano

provided in the fuselage to support NLG, wing, MLG, VT, HT and engine components which are all flanged I sections with stiffened web. Intermediate bulkheads are of one sided flanged sections (channel) distributed along the fuselage span to divide skin panels. Booms are not connected with the main fuselage beyond engine rear mount bulkhead. Specific areas like cockpit well, MLG cutout and engine door area schemes are studied. Highly stressed area in the engine duct will be minimized by the introduction of shear web between fuselage skin and engine duct. Finite element models are built separately for each of the fuselage segments (front, center & rear fuselage) and then integrated to have full fuselage finite element model. Finite element models are built using the tool ‘Altair-Hypermesh’ which is finite element modeling & analysis results post processing software. Static stress analysis of the entire fuselage for the specified critical flight load cases is carried out using MSC/Nastran which is a Finite Element Analysis (FEA) solver. Initially, fuselage has been sized based on the experience, design considerations and manufacturability issues. Structural (size) optimization of the fuselage for the specified critical flight load cases is carried out with an objective of mass minimization based on strength constraints. Layout of fuselage, finite element model and optimized thickness models are shown in Fig.21. The preliminary design, finite element analysis and optimisation studies of composite vertical tail and horizontal tail has been carried out for multi-spar construction and rudder multi-rib construction for four critical load cases (Mach 0.9 and 1.3, trimmed and untrimmed cases).The preliminary design, finite element analysis and optimisation studies of AMCA horizontal tail torque shaft has been carried out for the critical load case with 3 different materials. Vertical tail (VT) and Horizontal tail (HT) configuration, finite element models and horizontal tail torque shaft model are shown in Fig.22.

An approach for flutter prediction of aircraft structures in time domain using dynamic aeroelastic response analysis solver of NASTRAN The present work presents an approach for flutter prediction of aircraft structures in time domain using NASTRAN. Here, the dynamic aeroelastic response module of MSC Nastran is used to predict the flutter velocity of a swept back cantilevered plate in different flow regime. The finite element structural analysis of plate is performed using Nastran and the dynamic characteristics are compared with the experimental results. The aerodynamic model of the plate is done using doublet lattice method to predict the aerodynamic loads over the plate. The flutter analysis is carried out using two approaches: (i) Applying a discrete gust over the plate and (ii) Applying an impulse load at the tip of plate (Fig.23). Both the analyses are carried out at different velocities under sea-level conditions and the time histories of different responses at the tip of the plate are studied. The flutter velocity of the plate is then identified from the obtained responses.

Flutter prediction of swept wing like structures in time domain using fluid

structure coupling through 3D panel method

The present work deals with the aeroelastic behaviour of swept wing like structures. Here, a typical swept wing with arbitrary cross section is considered and modeled as a cantilever beam. A general Timoshenko beam theory is used to model the beam which takes into account both transverse shear effects and dynamical coupling between bending and torsion due to the fact that mass centre and shear centre do not coincide. The aerodynamic loads over the wing for various deformed positions are calculated using doublet lattice based 3D panel method. A FORTRAN code is developed to couple the aerodynamic and structural dynamic solvers in time domain. The coupled fluid structure solver computes the aeroelastic

Fig.20: Fuselage layout with Internal Members

Fig.21a: Finite Element Model of Fuselage (without skin)

Figs.21b: Fuselage Optimum Element Thickness (mm) Contour Plot

Figs.22: AMCA a) Vertical Tail, b) H T loads and boundary condition c) Sectional view and FE Model of H T d) Horizontal tail torque shaft

(a) V∞ = 305.7 m/s (b) V∞ = 305.71 m/s (c) V∞ = 305.73 m/s

Fig.23: Time history of wing tip responses due to the impulse load applied at the tip plate at

different flow velocities

response in the subsonic flow regime, using Newmark’s algorithm with suitable initial conditions. The results of free vibration and aeroelastic response of wing are also compared with the closed form solutions. The effect of bending rigidity and torsional rigidity on the frequency and flutter speed of the wing are studied. Further, the effect of sweep angle on the flutter speed of wing is also discussed.

-0.00010

-0.00005

0

0.00005

0.00010

0 0.2 0.4 0.6 0.8 1.0

Time (s)

Ver

tica

l D

isp

lace

men

t, w

(m

)

-0.00008

-0.00004

0

0.00004

0.00008

0 0.2 0.4 0.6 0.8 1.0

Time (s)

Vert

ical

Dis

pla

cem

en

t, w

(m

)

-5.0x10-5

-2.5x10-5

0

2.5x10-5

5.0x10-5

0 0.2 0.4 0.6 0.8 1.0

Time (s)

Ver

tica

l D

isp

lace

men

t, w

(m

)

Steady flow analysis over vertical fins of AMCA aircraft in transonic regime using Euler based CFD solver

The aim of this work is to perform steady flow analysis over twin vertical fins of AMCA aircraft in the transonic regime using the high fidelity CFD solver. Here, the flow analysis is carried out using inviscid flow assumptions at a free stream Mach number of 0.9. Various parameters such as pressure and Mach distributions on the surface of fin and fuselage and Cp distribution at various sections of fin are studied. Further, the steady loads on the fin are also evaluated. It can be observed that a large inward force Fy acts on the fin due to the presence of the tunnel (or passage) formed by both the vertical fins. Since there exists a shock near the rear end of the fin-fuselage junction (that has a wedge shaped structure), a new geometric model is generated by filling the wedge portion. Similar steady flow analysis is carried out for the newly generated model. The results indicate that the location of the shock on the fin surface remains the same. Thus it can be concluded that the location of shock is not due to the presence of wedge that exists in the original design.

(a) CFD Mesh (b) Pressure contours (c) Mach contours

Fig. 24a: Pressure and Mach contours on the fin and fuselage surfaces of AMCA aircraft

(a) fin root section (b) fin mid section (c) fin tip section

Fig. 24b: Comparison of Cp distribution at the root, mid and tip sections of the fin with two different meshes

Fluid structure interaction analysis of HIRENASD wing in time domain using coupled CSD-CFD solver

The aim of this work is to establish a computational procedure to perform the FSI analysis of the High Reynolds Number Aero-Structural Dynamics (HIRENASD) wing using the high fidelity CFD solver (Fig.25). First, the structural dynamic analysis of the wing is performed using the ABAQUS structural solver and the results of frequencies and mode shapes are validated with the experimental results. Then, the steady flow analysis over the wing and FSI analysis of flexible wing under steady flow is carried out at M = 0.8 and Re = 23.5 million. It is found that when compared to the steady flow analysis results, static FSI analysis results are very close to the experimental results. Further, the unsteady flow analysis over wing under forced oscillation with various modes is performed to get the time histories of force

-2.5

-2.3

-2.1

-1.9

-1.7

-1.5

-1.3

-1.1

-0.9

-0.7

-0.5

13.5 14 14.5 15 15.5 16 16.5 17

Press

ure C

oeff

icie

nt

(Cp)

Location (m)

Mesh - I

Mesh - II

-2.5

-2.3

-2.1

-1.9

-1.7

-1.5

-1.3

-1.1

-0.9

-0.7

-0.5

14 14.5 15 15.5 16 16.5 17

Press

ure C

oeff

icie

nt

(Cp)

Location (m)

Mesh - I

Mesh - II

-2.5

-2.3

-2.1

-1.9

-1.7

-1.5

-1.3

-1.1

-0.9

-0.7

-0.5

14.5 15 15.5 16 16.5

Press

ure C

oeff

icie

nt

(Cp)

Location (m)

Mesh - I

Mesh - II

and moment coefficients at M = 0.8 and Re = 23.5 million. This is followed by the dynamic FSI analysis of flexible wing. The nature of dynamic response is found to be converging indicating a stable system. This may be due to the fact that the present wing material has very high stiffness values (close to steel) which results in large difference between the bending and torsion frequencies.

(a) CFD Mesh (b) Pressure contours (c) density contours

Fig. 25a: Distribution of pressure and density over HIRENASD rigid wing for α = 0° with M

= 0.8, Re = 23.5 x 106 and q/E = 0.48 x 10-6

(a) wing root section (b) wing mid section (c) wing tip section

Fig. 25b: Comparison of pressure coefficients between rigid and flexible HIRENASD wing with experiments at various span wise stations for α = 0.0°, M = 0.8, Re = 23.5 x 106 and q/E = 0.48 x 10-6

(a) displacements (b) force coefficients

Fig. 25c: Time histories of displacements and force coefficients for the HIRENASD wing at M∞ = 0.8, ρ∞ = 2.8 and Vf = 0.0302

Structural layout and size optimization studies on a trapezoidal wing box like structure Structural optimization is an essential part of aircraft structural design and multidisciplinary design optimization process. The objective of this work is to setup a process that enables the automation of the structural layout and size optimization of wing box like structures. For this purpose, a trapezoidal wing box like structure is considered as a candidate for optimization satisfying the strength and buckling criteria, to arrive at an optimum structural layout (rib

-1.0

-0.5

0

0.5

1.00 0.2 0.4 0.6 0.8 1.0

ExperimentAECSteady-Rigid

x/c

Cp

-1.0

-0.5

0

0.5

1.00 0.2 0.4 0.6 0.8 1.0

ExperimentAECSteady-Rigid

x/c

Cp

-1.0

-0.5

0

0.5

1.00 0.2 0.4 0.6 0.8 1.0

ExperimentAECSteady-Rigid

x/c

Cp

-0.001

0.001

0.003

0.005

0.007

0.009

0.011

0 0.1 0.2 0.3 0.4 0.5

Tip Leading EdgeTip Trailing Edge

Time (s)

Win

g T

ip D

isp

lace

men

t (m

)

-0.16

-0.12

-0.08

-0.04

0

0.04

0.08

0 0.1 0.2 0.3 0.4 0.5

CL

CM

Time (s)

Fo

rce

and

Mo

men

t C

oef

fici

ents

positions) and sizing (skin/panel thicknesses) by minimization of structural mass or maximization of buckling load. Two optimization process flows are developed for the structural layout and size optimization of multi-rib trapezoidal box configuration subjected to elliptical load distribution. One of the optimization process flows is shown in Fig.26. The number of ribs are varied from one to five in a box with inter-rib panel aspect ratio (a/b) > 1.0 and from one to ten in the box with inter-rib panel aspect ratio (a/b) < 1.0. To implement the optimization process flows, CAD and FEA software are integrated with the help of process integration and design optimization software (Isight), which enables the automated and seamless work flow between CAD and FE modules. Mass minimization as an objective function is very effective while carrying out sizing optimization and the same is not true for layout optimization as mass is not sensitive to the rib position for a wing box with given number of ribs. Nonetheless, it is observed that for the wing box with inter-rib panel aspect ratio (a/b) < 1.0, variation in rib position played somewhat feeble role in enhancing the buckling load. However, the combination of rib positions and inter-rib panel thicknesses as design variables is the most effective in both the optimization cases i.e. mass minimization as well as buckling load maximization of trapezoidal wing box like structures.

Fig.26: Structural layout and size optimization process of trapezoidal wing box like structure

Demonstration of simulation data management in SLM: Structural optimization framework using Isight Apart from automating the simulation processes, the simulation scenarios and data needs to be securely stored and should be traceable and accessible. This could be possible with the software like SIMULIA SLM combined with ENOVIA technology. The present study is meant for understanding and utilizing the Dassault Systems (ENOVIA, SIMULIA) team work on SLM-Isight integration at NAL for Proof of Concept (POC). It may be noted here that this study is carried out with the support from Technical Specialist-SLM, Dassault Systems, India. The objective of this study is to demonstrate the simulation data management of ‘structural optimization framework using Isight’ in Simulation Lifecycle Management (SLM) of SIMULIA. In this regard, structural layout and size optimization of trapezoidal box structure is considered as a case study for simulation data management. Parametric CAD model generation (CATIA), automatic finite element (FE) mesh generation script file (HyperMesh) for FE model generation with analysis deck and finite element analyzer (MSC Nastran) are integrated using Isight for the optimization framework of structural layout and size optimization of trapezoidal box structure. The above optimization framework in Isight (sim-flow) is modified by adding ‘XML parser’ and ‘Data exchangers’, which transfer the input and output parameter values between Isight and SLM, Fig. 27a. The modified Isight sim-flow is

uploaded along with required input data to SLM and then a template is created for Isight based optimization framework in SLM. By using this template, simulations are carried out for Isight based optimization framework in SLM leading to structural optimization of trapezoidal box structure. And, the result files, consisting of Isight design cycles information along with optimized configuration of trapezoidal box structure’s CAD model, FE model and FE analysis result, are sent to the SLM, Fig. 27b. Hence, this exercise demonstrates the capability of SLM of SIMULIA for the simulation data management of Isight based structural optimization frameworks.

Fig. 27a: Structural optimization framework of trapezoidal box structure using Isight for SLM

Fig. 27b: Simulation data (input and results) of structural optimization

framework of trapezoidal box structure in SLM

Fatigue life prediction of plate with multiple stress concentration zones using

FEM

Present study discusses the available analytical fatigue life prediction procedures and their application to estimate the fatigue life of plates having single and multiple stress concentration zones without and with pre-existing cracks. First, finite element stress analysis of plates with single and multiple (three) circular holes is carried out to identify the stress concentration zones. Then, by taking the finite element stress analysis response as input, stress-life/strain-life approaches are used to estimate the fatigue life of plates having holes without pre-existing cracks. Next, fatigue crack propagation life of plates having holes with pre-existing cracks is estimated using crack growth laws in conjunction with finite element based fracture mechanics analysis. For example, in a plate having three holes arranged perpendicular to the load direction and fixed geometry (Fig.28a), the stress concentration will increase significantly when compared to single hole stress concentration and consequently fatigue life of the plate (without pre-existing crack) decreases significantly. Further, in a plate having three holes arranged perpendicular to the load direction and pre-existing crack, cumulative fatigue crack propagation life is estimated in progressive manner (Fig.28b) i.e. monitoring the crack propagation from inner-hole-edge to outer-hole-inner-edge and then from outer-hole-outer-edge to plate edge.

Product Lifecycle Management (PLM)

In modern aircraft design, development, manufacture and support, the engineering processes are to be managed under strict configuration control with model based definition and on concurrent engineering basis, covering parallel processing of design, prototyping, process design and maintenance design in order to reduce time to market on one hand and ensure that all aspects are taken into account early in design cycle. This is possible only by an Integrated Digital Environment. A highly interoperable CAD/CAE/CIM system called End to End PLM is organized to enable wide collaboration between widely separated engineering communities using diverse computing systems. The system ensures true concurrency. Further the practice of loft and physical mock up based aircraft design is to be replaced by the more efficient Digital Mock Up (DMU) based engineering, where all design engineers will add their portion of the product design on a visual DMU in Virtual reality, under configuration control. The DMU itself is created in a CAD/CAE/CIM neutral, light weight format to enable diverse Engineering groups to work and build the aircraft in virtual environment and thus reduce reengineering, development time and cost. There are many PLM solutions that can be added on to the core technology of PLM. There are many commercial PLM software tools available today and few of them being used by Aerospace industries across the world for their aircraft development programmes.

As part of the PLM software tool evaluation process at the division, pilot project is initiated with Dassault Systèmes (DS) using the tool ENOVIA which is fully integrated with design, engineering and analysis solutions from Dassault Systèms and other software vendors.

Fig. 28a: Plate with three holes arranged

perpendicular to load direction. Fig. 28b: Fatigue crack propagation life of

plate with three holes (R = 0.1)

Team of scientists from the division has been formed to take part in the software evaluation process. Typical procedures being followed covering various processes like customer enquiry, project proposal and project execution with tight integration of CAD and CAE tools have been proposed for the pilot project. Various CAD and CAE tools (CATIA, Hypermesh, MSC/Nastran, Isight) being practiced are considered to be part of the evaluation process to verify the integration with the PLM tool. DS team is carrying out the pilot project at the division by having interaction with the team identified from the division and this will be followed by demonstration of the same illustrating the capabilities and usage of the tool ENOVIA.

Studies on embedded delamination and debond in stiffened composite panels

Analytical solution for buckling load of a laminated composite plate containing embedded delamination is determined. The plate is subjected to uniaxial uniform compressive load and clamped on the loading edges. The effect of crack size and fiber orientation on the buckling load is determined analytically. Using the principle of minimum elastic potential of the plate in the buckled state and boundary & continuity conditions, an explicit representation for the buckling load is determined in a delaminated plate. Using Von-Karman non-linear strain-displacement relations and Ritz method an analytical methodology for the calculation of strain energy release rate is developed and SERR is calculated for a composite delaminated plate. Thickness dependent stress intensity factor in an isotropic plate containing crack determined analytically and verified with finite element results.

1E. Mechanical Systems Design Group

Ejector Pump Performance Assessment

Ejector pumps are secondary boost pumps in the aircraft fuel system.They maintain 100% supply at boost or primary pump inlet (FAA Chapter 14) and scavenges the fuel from the remote locations of the tank. In small aircrafts like CRJ 200, ejectors are primary boost pumps no electric or mechanical boost pumps exist. Ejectors are more reliable and safer in operation as there are no moving parts. Major disadvantage of ejector pumps is low in efficiency (25-30%). To analyze the performance of ejector a test rig with scheme shown in Fig.29 is proposed and the rig is under commisioning. Simulation model will be built for the ejector and will be benchmarked with test results and further design studies will be carried on the simulation model

Fig.29a: Test rig schematic of Ejector pump 29b: Ejector Test rig

Cabin Pressure Control System for Transport Aircraft

Cabin pressure control system works in conjunction with environmental control system to maintain and regulate pressure within the aircraft cabin as per design intent. A study was carried out to identify the system architecture, design considerations, weight estimation, system interfaces, testing, integration and development methodology as well as safety and reliability aspects to be considered within the framework of FAR-25 regulations to be employed while deploying a transport aircraft.

Study of Flapping Wing Mechanisms for Ornithopter Applications

Synthesis of flapping wing mechanism (ornithopters) for generating a figure-of-eight path similar to humming bird morphology is a challenging task. A study has been carried out to determine the potential mechanisms that generate figure-of-eight path. A novel spatial mechanism based on simple planar four bar for generating spatial figure-of-eight path has been synthesized and analyzed. A planar four-bar mechanism can be used for generating various shapes of coupler curves depending upon the linkages configuration. The figure-of-eight coupler curve is of particular interest in the design of flapping wing mechanism for ornithopter. A sensitivity study has been carried out to determine the effect of various parameters of the four-bar mechanism on the f-o-8 coupler curve. A program has been developed to synthesize a four-bar mechanism based on several design requirements, important being the minimum transmission angle and mechanism bounding box requirements. A kinematic and dynamic analysis program has also been developed to determine the velocity, acceleration, joint forces and torque required to drive the mechanism. The results from the analysis were compared and validated using the MSC/ADAMS model results.

Pressure loss estimation in the Bleed Flow Pipe line connecting between Auxiliary Power Unit (APU) and Air turbine starter (ATS) of an Aircraft Air Turbine engine Starter (ATS) runs with the high pressure air bled from Auxiliary Power Unit (APU) compressor. The bleed air is supplied to ATS inlet through a pipe layout, which leads to a permanent pressure loss due to pipe friction and angular bends. After the loss, the pressure of bleed air should be always above the minimum required pressure of ATS inlet. Pressure loss due to pipe friction and angular bends is estimated for a particular layout of AMCA aircraft using both compressible and incompressible flow equations. Even though the flow is in incompressible regime, pressure loss is estimated with both incompressible and compressible flow equations to establish the methodology and results comparison. The methodology to estimate bleed flow parameters at altitude condition is discussed in detail. Apart from ground start, Military aircrafts will be restarted at certain altitudes as well in case of any failure. The bleed pipe layout should be designed for minimum pressure losses at altitude condition as well, which requires bleed flow conditions to be known at that particular altitude (Fig.30).

Fig.30a: APU to ATS Bleed flow pipe Fig.30b: Incompressible and compressible layout flow results comparison for the given pipe layout

2. ANALYSIS AND DESIGN ACTIVITIES 2A. Adaptive Aeroelastic Structures and Vibroacoustics Vibroacoustic modeling / testing of fuselage under low, mid and high frequencies with active/ passive solutions In the present work, vibroacoustic modelling methodology for the simulation of noise transmission into aircraft cabin of a segmented fuselage of 5m length is being developed. The finite element model of the segmented fuselage structure, Fig.31 is made as a stiffened structure; which includes structural members like bulkheads, stringers that add strength to the aluminium skin (outer/inner) and the windows are modelled with acrylic material. Floor board and luggage board is also considered in the model. Static analysis is carried out to check the strength of the model by applying an internal cabin pressure of 12 psi and the stresses are observed to be within the allowable limits. The cabin air is modelled as fluid element with atmospheric properties. Evaluation of Fiber Metal Laminate (FML) panel’s response under acoustic loadings, its potential in sound transmission and effects of delamination on vibroacoustic performance is

evaluated. For this a FEM based impedance tube model is developed in ANSYS and vibroacoustic analysis is subsequently carried out on aircraft panels. The numerical impedance model, Fig.32, has been found suitable to simulate the VA behaviour of panels made of different class of materials. From the study it’s found that the presence of delamination in laminated/FML structures has effected the Sound Transmission Loss (STL) and the local delamination modes influence the transmitted noise to get distributed with nearby global modes, around its local frequency, Fig.33.

10-50 50-100 100-150 150-200 200-250 10-250

0

20

40

60

80

100

OA

SP

L i

n d

B

Frequency in Hz

AL

Composite

FML

Fig.32: Finite element model of the digital

impedance tube Fig.31: Finite element model of

segmented fuselage

Fig.33: OASPL of different aircraft materials

Vibro-acoustic response and sound transmission loss analysis of functionally

graded plates

Functionally graded materials are a class of composite materials in which the material properties vary continuously from one surface to another to achieve a desired functional performance. Analytical studies were carried out on the vibro-acoustic and sound transmission loss characteristics of Functionally Graded Material (FGM) plates using a simple first order shear deformation theory. The material properties of the plate are assumed to vary according to power law distribution of the constituent materials in terms of volume fraction. The sound radiation due to sinusoidally varying point load, uniformly distributed load and obliquely incident sound wave is computed by solving the Rayleigh integral with a primitive numerical scheme. Displacement, velocity, acceleration, radiated sound power level, radiated sound pressure level and radiation efficiency of FGM plate for varying power law index are examined. The sound transmission loss of the FGM plate for several incidence angles and varying power law index is studied in detail. It has been found that, the sound power level increases monotonically with increase in power law index at lower frequency range and a non-monotonic trend is appeared towards higher frequencies as shown in Fig.34. Increased vibration and acoustic response is observed for ceramic-rich FGM plate at higher frequency band; whereas a similar trend is seen for metal-rich FGM plate at lower frequency band. The vibrations of lower order modes have made a significant contribution to the radiated overall sound power in the complete frequency band. The radiation efficiency of ceramic-rich FGM plate is noticed to be higher than that of metal and metal-rich FGM plates. The transmission loss below the first resonance frequency is high for ceramic-rich FGM plate and low for metal-rich FGM plate and further depends on the specific material property. A typical STL plot of FGM plate is shown in Fig.35. The study has found that increased transmission loss can be achieved at higher frequencies with metal-rich FGM plates. Several practical applications are related to this formulation and study, specific to aerospace and automobile industry where more usage of FGM plates is envisaged. It is possible to analyze and optimize the vibro-acoustic behaviour of panels as an integral part of the design process for improved sound transmission characteristics. The other applications of this study are related to parameter optimization, sensitivity analysis and vibro-acoustic system identification.

Fig.34: Sound power radiated by FGM plate under centrally Fig.35: Typical sound transmission acting unit mechanical force loss plot of FGM

Multifunctional Composites

Structural Health Monitoring

Design and Development of Numerical and Experimental Schemes for Structural

Health Monitoring

Here we address the structural health monitoring through development of numerical and experimental schemes. A 36-noded piezoelectric spectral plate element has been designed

and developed to model plate structures with surface-mounted piezoelectric transducers. The developed element independently captures the kinematics of the layers of piezoelectric, bonding and host structure (plate). The layer-wise kinematics is then coupled through Lagrange multipliers and the element thus formulated has a diagonal mass matrix, which is a basic requirement of any time integration scheme. Fig.36 shows the 36-noded piezoelectric spectral element and Fig.37 shows the simulation of wave propagation through the plate structure. Further, experimental schemes are developed to simultaneously identify the transducer debond and structural damage. Structural damage shows a small increase in damage index gradient, whereas larger reduction in damage index is observed when debond occurs in the sensor. Fig.38 shows test set-up used for conducting the Lamb wave experiment, with eight sensors distributed in the circular pattern and has different debond types. The damage index, where the gradient change due to structural damage and sensor debond are explicitly shown in Fig.39.

Fig.38 : 36 noded spectral element Fig.37: Propagating waves in a plate

Fig.38: Experimental test set-up Fig.39: Damage index due to sensor debond

and structural damage

Finite element modeling of intra-plate in peninsular India The orientation of maximum horizontal compressive stresses in the earth’s crust gives the insight into the driving mechanisms of plate motion and intra-plate seismicity. In an attempt to numerically model the maximum horizontal compressive stress distribution in peninsular India, the Indo-Australian tectonic plate has been modelled by considering homogenous material distribution. The benchmarking and validating of the procedure has been done on Australian continent, for which the stress patterns are available in literatures. The finite element model of the Indo-Australian plate with boundary conditions and external loading, the displacement pattern of the plate, and the maximum horizontal compressive stress directions in the Australian continent are shown in Figs.40 - 41, respectively. Further, the compressive stress distribution in peninsular India, considering homogenous and heterogeneous material properties on the plate is planned.

Fig. 40: CAD and Finite element model of the Indo-Australian tectonic plate with boundary

conditions and external forces

Fig. 41: Displacement pattern of the Indo-Australian tectonic plate

Maximum horizontal stress directions in Australian continent

2B: Whirl Flutter Analysis of Pusher Type Propeller- Necelle System of SARAS

PT1N using a 2 DOF Model

The objective of this work is to study the whirl flutter characteristics of SARAS PT1N engine propeller system using an analytical procedure based on two-degree of freedom model (pitch and yaw) (Fig.42). Here, the whirl flutter analysis is done using the data (geometric/inertia/aerodynamic etc.) supplied by C-CADD and Ground Vibration Tests (GVT) data given by GVT team. The experimental modal parameters used in the whirl flutter analysis are the natural frequencies, mode shapes and modal damping for the engine-propeller system. Results of whirl flutter analysis of SARAS PT1N engine-propeller system are computed for two flight conditions, namely, sea level and 15000 ft. At each altitude, the forward and backward whirl frequencies and the corresponding damping necessary for stability boundary conditions are generated for different design speeds. From the results it is noticed that the theoretical pitch damping is negative for all airspeed upto 195 m/s. It indicates that the critical whirl flutter velocities of the engine-propeller system are much higher than 1.2 VD at all the altitudes considered. A report has been published in the form of NAL project document.

(a) Sea level (b) 15000 ft

Fig. 42: Variation of pitch damping ( g ) of SARAS PT1N engine propeller system for different

airspeed at sea level and 15000 ft

Reliability Analyses for SARAS PT1N

The PT1 aircraft being the first prototype aircraft of SARAS, the FHA and FTA were carried out for all the systems. In order to overcome some of the deficiencies observed in PT1 aircraft, the aircraft has been modified in respect of engine, propeller, nacelle, stub wing, flap-tracks, structural elements (VT, HT), U/C, Brake Management System, Environmental Control System, Rudder, Flight Control System (FCS) etc and renamed as PT1N aircraft. The System Safety and Reliability studies having been carried out both for PT1 and PT2 aircraft, the changes in terms of modifications carried out on PT1N aircraft have now been analyzed for the impact and cascading effects on safety. In view of the changes, it was considered fit to carry out the system safety assessment and FTA of the FCS for the identified critical events from FHA, which are probable for PT1N aircraft and its compliance to the safety requirements laid down as per FAR-23. Accordingly the FHA and FTA of the flight control system of SARAS PT1N have been addressed. The Failure Modes & Effects Analysis (FMEA) is carried out as a part of safety assessment requirement as per ARP[1] provides an in-depth analysis at each LRU level with a view to identify potential failures and their effects on the equipment, sub-system and at aircraft level. FMEA of the flight control system has been carried out, Fig.43.

Fig.43: Schematic of Aileron Roll Control.

-0.08

-0.06

-0.04

-0.02

0

0.02

0.04

0 50 100 150 200 250

dam

pin

g (g

θ)

Velocity (m/s)

Forw…Back…

-0.06

-0.04

-0.02

0

0.02

0.04

0 50 100 150 200 250

dam

pin

g (g

θ)

Velocity (m/s)

Forw…Back…

The SARAS PT1N aircraft is an update of SARAS PT1 aircraft, wherein one of the major modifications is the engine change from PT6A-66 to higher rated PT6A-67A engine of M/s Pratt & Whitney, with a higher dimension propeller. To cater for the higher rated engine, the mounting assembly has been augmented with additional trusses to interconnect yoke with L and M frames. Thus, it was considered necessary to revisit for the identified critical events of the FHA and to carry out the FTA for the engine and its interface systems on aircraft. FMEA of the engine-aircraft interfaces has been carried out. The Zonal Safety Analysis (ZSA) is one of the essential activities in the system safety assessment process for any aircraft development programme. Despite layout studies and engineering mock up studies, physical installation of system hardware as a whole is the one addressed in the ZSA. It is therefore necessary that the system configuration for safety, independency, absence of interference, case of maintenance and supportability are looked for in detail when the system matures. The ZSA is invariably carried out during the integration and installation phase. In view of the large scale changes carried out on PT1N aircraft, it was considered fit to revisit ZSA as part of the system safety assessment. The electrical system in SARAS aircraft plays a vital role for the functions of both power generation and distribution. The PT1-N architecture is a partial update over the PT1. Though the FHA for SARAS electrical system has been carried out earlier, in view of certain additional elements being added the FHA has been revisited and the undesired events have been identified, besides the FTA for the identified critical events. FMEA of the electrical system has been carried.

Reliability Analyses for systems of AMCA.

The Loss of control (LOC) of Integrated Flight Control System (IFCS) is a catastrophic event and it is customary to predict the quantitative probability (PLOC) and to derive the redundancy levels. In order to maximize the performance, for the dictated PLOC as per regulatory requirements, various options in terms of redundancy (for two cases viz. quadruplex and triplex architecture) have been evaluated. The failure probabilities of aircraft electrical and hydraulic systems have been obtained through an exclusive fault tree analysis and used in the Probability Loss of Control (PLOC) calculations. The use of COTS technology components have also been plugged in and the PLOC revisited for the triplex architecture, having in-line monitoring provision, while there is no other change in respect of rest of the interfacing sub systems, Fig.44.

Fig.44 : PLOC for AMCA (Mil-Based) 2/4 Quadruplex Architecture

The Failure Modes and Effects Analysis (FMEA) for the flight control system have been carried out for the cases where it would affect the sub-systems and its functions on aircraft, considering all possible failure modes of the components in the flight control systems. The Failure Modes and Effects Analysis (FMEA) for the Electrical System has been carried out. The Failure Modes and Effects Analysis (FMEA) for the hydraulic system have been carried out for the cases where it would affect the sub-systems and its functions on aircraft, considering all possible failure modes of the components in the hydraulic systems. The Failure Modes and Effects Analysis (FMEA) of the proposed Secondary Power System (SPS) for Advanced Medium Combat Aircraft (AMCA)has been carried out. Additionally, in order to identify the elements for design improvements, quantitative analysis to determine Risk Priority Number (RPN) has been carried out. A cascading failure, which is a “low-probability high-consequence event”, is a particular type of common-mode failure in which a single event, not necessarily hazardous in itself, can precipitate a series of other failures, which may endanger either the mission or safety of the aircraft. In the initial stages of the program, the cascading failure analysis has been carried out qualitatively for the inter-dependant systems through Cause-Mode Effect Analysis. Dynamic and Flutter Analysis of SARAS PT1N Aircraft.

The dynamic and flutter characteristics for SARAS PT1N aircraft has been studied using MSC/NASTRAN for two configurations namely the takeoff (30.4% MAC, 6700 kg) and landing (29.6%MAC, 6206 kg) configurations. Doublet lattice method has been used to estimate the unsteady air loads on the lifting surfaces and flutter analysis has been carried out by P-K method. Flutter velocities and margins have been established for both the configurations using the design control circuit stiffness and also by fine tuning the control surface frequencies as obtained from GVT (Fig.45). Parametric studies were carried out to study the effect of variation of the control circuit stiffness and mass balancing on the flutter speed of aircraft. Additionally the flutter speeds have also been estimated using the experimental modal parameters obtained from ground vibration tests and using doublet lattice aerodynamics using NASTRAN. The detail analysis with the flutter speeds and margins have been reported.

.

Fig.45: Structural and Aerodynamic Mesh of SARAS PT1N.

Flutter Analysis of an Utility Aircraft using Experimental Modal Parameters

and ZAERO

The flutter characteristics of an Utility aircraft for full fuel and empty fuel configuration has been studied using ZAERO and experimentally obtained modal parameters. ZONA6 aerodynamics has been used to estimate the unsteady air loads on the lifting surfaces and flutter analysis has been carried out by g-method (Fig.46). The detail analysis with the flutter speeds and margins have been reported.

Fig.46 : Aerodynamic Model of Aircraft.

Flutter Analysis of a Typical Combat Aircraft using Experimental Modal

Parameters.

Flutter results for a typical light combat aircraft using GVT results have been evaluated using equivalent finite element structural models in NASTRAN and experimental modal parameters from GVT. Doublet lattice and ZONA51 aerodynamics have been used for carrying out aero elastic flutter analysis. The flutter characteristics are evaluated for flight conditions, namely at sea level, 5 km, 10 km and 15 km. The results are compared with those obtained from flight flutter testing. The aircraft is free from flutter upto a Mach no. of 2.0 at altitudes of 10 km and 15 km.It has been noticed that the aircraft is free from flutter in subsonic regime and at altitudes of 10 km and 15 km in supersonic regime, Fig.47.

Fig.47: Aerodynamic Model of Combat Aircraft.

Reliability analysis by Parallel Monte Carlo simulations using OpenMDAO

Reliability analysis is of significance as far as safe operation of an aircraft is concerned. Structural reliability analysis quantifies the effects of uncertainties in the structure, which will help in design of new structures as well as maintenance of the structures in service. Reliability analyses on a benchmark wing structure and truss from literature and a wing-box segment are done using Monte Carlo simulation. A python based software platform called

Open-MDAO has been used to generate randomized uncertain structural parameters to update Nastran solver deck and set up Monte Carlo simulations. The probability of failure has been determined based on the defined limit states for each structure. Each analysis in Open-MDAO is configured to run the simulations in parallel to make use of multiple processors.

Static Aeroelastic Loads Analysis of Advanced Medium Combat Aircraft (AMCA)

This work describes the process of computation of the load distribution, on the aircraft, when it is performing any particular manoeuvre. The load distribution considers the influence of the aerodynamic pressures at the specified flight condition, the inertial parameters (due to a specified mass distribution) and also the modification of the effective load distribution due to aeroelasticity. The entire process of load computation has been established within the Finite Element software, MSC/NASTRAN. The aerodynamic pressure distributions, for the present exercise, have been computed using the Euler Based CFD Code, MGAERO. The initial FE model, considered for this set of load computations, has been arrived at based on certain preliminary calculations, empirical relations and experience. The ‘Rigid Manoeuvre Loads’ obtained (aero elasticity not considered) have been transferred to the Design Groups for a more detailed Sizing and Stressing exercise. At the end of this exercise, a more accurately sized aircraft structural model is available. The load computation also describes the computation of the Aeroelastic Efficiencies, associated with the various aerodynamic effects. The more accurately sized model, referred above, is used for these computations. The efficiency parameter is a measure of the stiffness of the aircraft, Fig.48.

Fig.48: Finite Element Model and Aeromesh

Optimization of isotropic cylindrical shells using semi analytical approach for

gradients

In the present work, single objective optimization of isotropic shell structures are studied using semi analytical gradient based Sequential Quadratic Programming (SQP) approach. First, the structural static and dynamic analyses are carried out using a four noded shell element based on degenerated shell theory. The static and dynamic sensitivity analyses are performed using both Semi-Analytical (SA) and Finite Difference (FD) methods by treating the thickness of shell as design variables. Both uniform and strip wise variations in thickness are considered in the formulation. A special attention is given on the accuracy of sensitivity calculations and performances of each method. Further, the optimization problem is formulated as single objective, to minimize the weight of the shell with uniform and strip wise

variations in the thickness subjected to maximum displacement and natural frequency constraints. The optimal solutions are obtained by integrating the finite element method, sensitivity analysis and sequential quadratic programming method. Numerical studies are carried out to get the optimal thickness for shells under various boundary conditions. The effect of uniform, strip wise variation in thickness on the optimum design of shells is also studied. A report has been published in the form of NAL technical memorandum /1/.

2C. Layout, Preliminary Landing Gear Design of the Advanced Medium Combat

Aircraft (AMCA)

Layout and Preliminary design of Main Landing Gear (MLG) has been carried out for the Advanced Medium Combat Aircraft (AMCA) as per MIL-A-8862 load requirements. (Fig.49). Positioning of the main landing gear with respect to centre of gravity of the aircraft, and the height of main landing gear have been determined from considerations of stability and ground clearance. Single stage oleo-pneumatic telescopic shock absorber has been adopted for absorbing dynamic loads (fig.50). The tyre selection has been carried out based on the induced static loads, dynamic loads and brake energies requirement. Extension-retraction kinematics has been designed to suit the bay volume requirements and sustain landing and ground loads. The attachment loads have been estimated using MSC/ADAMS software. Preliminary sizing of the landing gear components has been carried out based on critical Landing and ground handling cases. The required air volumes for the shock absorber has been estimated based on the maximum dynamic loads.

Fig.49: Details of MLG Spatial Kinematics

Design and Development of Ultra High Pressure System for Food Preservation

Technology

An ultra high pressure system for Food Technology Research Applications has been designed and developed by CSIR-NAL for Central Food Technological Research Institute (CSIR-CFTRI), Mysore. The Ultra High Pressure System consists of a water compatible 6900 bar (690 MPa) capacity ultra high pressure pumping unit and a 4000 bar (400 MPa) capacity water compatible high pressure chamber. The ultra high pressure water pumping unit is basically an air operated system. The detailed technical specifications of the pumping unit and the corresponding system level design are carried out by NAL.

0

100000

200000

300000

400000

500000

600000

0 50 100 150 200 250 300

Load

, N

Deflection, mm

Load Vs…

Fig.50: Air Curves for Shock Absorber

Design and Development of Ultra High Pressure System for Food Preservation

Technology

An ultra high pressure system for Food Technology Research Applications has been designed and developed by CSIR-NAL for Central Food Technological Research Institute (CSIR-CFTRI), Mysore. The Ultra High Pressure System consists of a water compatible 6900 bar (690 MPa) capacity ultra high pressure pumping unit and a 4000 bar (400 MPa) capacity water compatible high pressure chamber. The ultra high pressure water pumping unit is basically an air operated system. The detailed technical specifications of the pumping unit and the corresponding system level design are carried out by NAL.

The high pressure chamber consists of a threaded end cap and a mono block cylinder with a bore of 55 mm internal diameter with approximately 1 litre working volume. The material of the high pressure chamber is Maraging steel MDN-250 (Fig.51). Preliminary design of the high pressure chamber, detailed Strength of materials design, Finite Element Analysis, validation of the Chamber design with respect to ASME BPVC, detailed drawings, chamber manufacturing process specifications are carried out by NAL. The high pressure chamber along with the ultra high pressure pumping unit has been successfully installed and commissioned at CFTRI, Mysore. Inspection and testing of the high pressure system have been carried out and the operation and performance of the system are found to be satisfactory. The end users have conducted more than 100 successful high pressure studies on the food specimens for their research applications and the performance of the system is found to be satisfactory (Fig.52). Fig.51: High Pressure System Fig.52: High Pressure Chamber

Landing Gear Brake system mathematical modelling and simulation

Landing Gear Brake system Matlab/Simulink model has been developed to predict the stopping distance and stopping time of the aircraft upon brake application. The braking system controller regulates the skid servo valve current such that maximum deceleration of the aircraft is below 7 m/s2 for dry run way condition. The Simulink model calculates the aircraft speed based on the linear momentum balance of the entire aircraft. In the aircraft linear momentum balance, only inertial and friction forces are considered and aerodynamic drag is neglected. The wheel velocity at any given instant of time is calculated based on angular momentum balance of the wheel. Angular momentum balance equation constitutes applied brake torque and tyre frictional torque. Skid controller is modelled as on/off controller in Simulink (Fig.53a & b).

Fig.53a: Matlab/Simulink model of Landing Brake system

Environment Control System Design and Analysis Activities

Layout and Preliminary design of Environment Control System has been carried out for a Transport Aircraft configuration as per FAR-25 requirements. Studies were carried out with respect to design aspects, basis of selection among various systems, bleed air penalty comparison among various architectures, performance estimates for various design cases, weight and reliability estimates as well as the design, development, testing and integration methodology to be employed for deploying a transport aircraft. Heat load estimates of critical design cases have been carried out and layout has been prepared in line with aircrafts of similar category, Fig.54.

Fig.54a: Layout of ECS b: Air Cycle Machine

C: Isometric view of the spatial mechanism

Fig.53b: Calculated Vehicle speed

& wheel speed of aircraft

Vs Time

d: Power and Torque requirements of the Mechanism

Static Burst Analysis Of Thick Walled Unflawed Cylinder

A study has been carried out to determine the static burst pressure in a closed unflawed thick cylinder, Fig.55. The failure of the cylinder by ductile fracture mode has been investigated. An elasto-plastic analysis has been performed considering the geometric and material nonlinearities using commercially available ANSYS finite element analysis (FEA) code. The Newton-Raphson and Arc-length algorithms available in ANSYS code has been employed for the solution. In the initial analysis, a closed cylinder with a constant ratio of outer to inner diameter, K =3 has been chosen. The principal stress distribution across the cylinder wall in the elasto-plastic and full plastic regimes has been validated using Huang's model. The Arc-length algorithm has been employed for computing the static burst pressure of the cylinder (over the range K = 1.5 to 6) through rigorous computations. The results are compared with the numerous empirical burst models available in the literature over the above said range for various material models, viz., Elastic Perfectly Plastic, Elasto-Plastic and Non-Linear.

Axisymmetric segment with boundary conditions for FE-analysis & Von Mises contour plot

fully plastic regime of the cylinder wall

Fig.55: Comparison plot of burst pressure for non-linear model

Static limit load test on engine mount structure: SARAS-PT1N aircraft:

The engine mount structure of the SARAS PT1-N has been redesigned to have a fail-safe design to cater uncontained rotor burst requirement, for which the engine mount design has been modified with an introduction of fail-safe truss assembly between yoke assembly and L and M frames of nacelle. The newly designed nacelle is part of the engine mount structure to fulfill the requirement of sealing. It was required to qualify the newly designed engine mount of SARAS PT1-N aircraft by carrying out a static load test for a critical design limit load case namely Propeller Malfunction to clear for flights including first flight of SARAS-PT1N aircraft. Static load test rig on the stub wing engine mount assembly is shown in Fig.56. Loads on the stub-wing were applied using lead shots bags. Engine mount loads at four mounting points were applied through mechanical turn-buckle arrangements. Also, loads on the L and M frames were applied by turn buckle and pulley arrangements. The test article was strain gauged at critical locations using linear gauges (72 nos.) and rosettes (23 Nos.). Data acquisition system (System 5000 from Micro- measurement group USA) was used for acquiring the strain data from all the gauge locations. Deflections during the tests were measured using two Cable Displacement Sensors (CDS) at identified locations. The test was carried out successfully and the strains monitored during the test were well below the allowable limits. The test was witnessed by CEMILAC certification agency.

2D. Static structural strength evaluation of RUSTUM-II fuselage:

The Rustom-II is a Medium-Altitude, Long-Endurance (MALE) Unmanned Aerial Vehicle (UAV) with AUW of 1800 Kgs. It is proposed to carry out static strength test of Rustom-II fuselage for proof loads to meet the design and certification requirements. The structural loads considered for the structural testing of the fuselage include the inertial loads experienced by the UAV and manoeuvre loads on the tail of UAV. For this, test rigs and loading fixtures are designed and developed for carrying out the static strength test for two flight load cases and a Landing case. A conceptual test rig set-up and loading arrangements is shown in Fig. 57. Preparations for test set-up and loading arrangements are underway.

Fig 56: Static load test set up for engine mount

structure: SARAS-PT1N aircraft

Fig. 57: Conceptual test setup for structural strength evaluation of RUSTOM-II UAV fuselage

Damage Tolerance testing and evaluation of LCA wing-root fitting box:

As a continued activity of LCA airframe testing and analysis, preparations for fatigue testing of the LCA wing-root fitting box has been completed. The test box has been assembled. The top and bottom skins of the box were impacted at few specified locations with designated impact energies and then C-scanned. In order to carry out structural health monitoring during testing, FBG sensors and strain gages have been installed both inside and outside of the test box at critical locations. Fatigue testing under variable amplitude loads will be carried out in due course of time.

FE Analysis of CFC panels subjected to tensile loads:

The CFC AS4/914 panels of size 300mm ×1000mm with sixteen-ply and quasi-isotropic (45/-45/0/0/90/45/0/0) lay-up sequence were modeled. The repair of damaged panels by adhesive Redux 319A with scarf patching geometry were considered. All the three configurations namely the undamaged, damaged (with centre hole) and the repaired were analyzed for their strength in order to examine the effect of damage on the strength reduction in the damaged specimen and to assess the performance of the repaired panels in regaining the strength of the undamaged panel. Failure load for undamaged, damaged and repaired composite panels was predicted by finite element methods and compared with the experiments. The FEA failure loads were obtained based on first ply failure theory, using Tsai-Wu failure criteria. On comparison of these values with experimental failure loads, it was observed that failure load obtained by FEA were in very good agreement with experimental results. Also, the FEA strain values predicted at various critical locations in all the three panels were in good comparison with experimental strain values. A typical comparison of strains obtained during experimental study are compared with FE results is shown in Fig.58.

Fig. 58: A comparison of FE and experimental strains on the repaired panel

3. TESTING AND QUALIFICATION ACTIVITIES

3A. Ground Vibration Testing on SARAS PT1N

The Ground Vibration Testing (GVT) was successfully conducted on SARAS PT1N aircraft which is an equivalent version of PT2 with additional features and design improvements. The GVT is mandatory for a newly designed aircraft and also for an existing aircraft with design modifications involving structural changes. The GVT was carried out to demonstrate the compliance with the provision of FAR 25 dealing with the design requirements for transport category airplanes (FAR Advisory Circular number 25.629-1A dated 23/7/98). The objective of the GVT was to determine the dynamic characteristics of the SARAS PT1N aircraft in terms of experimental modal parameters. The experimental modal parameters are the natural frequencies, modeshapes and modal damping. The GVT was carried out for “Take-off” and “Landing” configurations as per the test plan TS-03, Vol 6(1), 2013, approved by CEMILAC. The free-free boundary condition was simulated by reducing the pressure of tyres to half of its nominal rating. Multi-Input Multi-Output (MIMO) methodology was adopted as test strategy for global mode identification and Single Input Multiple Output (SIMO) for control surface mode identification. The aircraft was excited using four 200 N B&K electro-dynamic shakers, two of them located at wing tips in vertical direction, one at HT tip in vertical direction and the other at the rear fuselage inclined at 45ºin XZ plane. A total of 160 numbers of accelerometers and 4 numbers of force transducers were used for output response and input force measurement, respectively. The instrumented aircraft is shown in Fig.59. The state-of-the-art LMS SCADAS III hardware with 4 Digital to Analog Converter (DAC) and 126 Analog to Digital Converter (ADC) channels was used as data acquisition system. LMS Test.Lab software was used for data collection, processing and estimation of modal parameters. A frequency band of 0-100 Hz with 256 spectral lines was chosen to cover the aircraft global modes and control surface modes. A 50 % burst random signal for excitation with 50 averages was used to have good signal to noise ratio in the measured signals. Pretest verifications like Linearity, reciprocity, coherence and drive point FRF were performed before the estimation of frequency response functions (FRF’s). The estimated FRF’s were processed to identify the modal parameters. Further, modal validation was done

in terms of Modal Assurance Criterion (MAC) matrix. The control surfaces (Rudder, Flaps, Ailerons and Elevator) modes which were not distinctly excited during global excitation were excited using an instrumented hammer and the corresponding modes were identified by SIMO approach. A rigid body mode (Roll) was observed in all the test configurations. The elastic modes of the aircraft were carefully analyzed and grouped as global and control surface modes. A typical aircraft modeshape is shown in Fig.60. The estimated modal frequencies and modal damping values showed consistency and expected trend among the configurations. A detailed SIMO test was also conducted on both the engines in order to estimate the pitch and yaw modes of the engine. All the above tests were witnessed by the certifying authority (CEMILAC).

Fig.59: Instrumented SARAS PT1N aircraft Fig. 60: Wing 1st Symmetric Bending Mode (6.33 Hz)

Use of Multifunctional Materials in MAV’s for Improved Aeroelastic Performances The objective is to develop multifunctional wing for MAV to achieve required trim conditions through adaptive trailing edge. In this project, we have successfully employed Macro Fiber Composite (MFC) actuators on NAL developed BK MAV to realize the hingeless control surface concept through a trailing edge morphing. Numerical analysis, followed by experimental validation have confirmed that the trailing edge morphing technique is useful for achieving the aerodynamic trim conditions for takeoff, level flight etc. The morphed surface can be deployed as elevators (symmetric) and or ailerons (anti-symmetric) in flight for different flight maneuvers (Fig.61).

Fig. 61: Multifunctional structural analysis and testing for trimming (upward and downward continuous camber change)

a. Energy Harvesting Systems

A series of experiments are conducted to optimize the energy harvesting design sensitivity

parameters (Fig.62). The experimental details are given below:

i. Estimation of energy dissipation from 40X20 PZT patch

ii. Study on the effect of the thickness of the sensor to the thickness of the structure

iii. Effect of the vibrating frequency on harvesting energy

iv. Energy harvesting using Micro Fiber composite (MFC) to understand the directional

effects on energy harvesting

v. Study on dimensional effects of PVDF films on energy harvesting

Fig. 62: Energy Harvesting System setup and results of the power output Vs applied force

Nanocomposites

CNT alignment in epoxy composites has been examined to improve the mechanical and damping properties. Voltages are applied during fabrication stage to electrically align the CNTs. Specimens are fabricated and tested in Dynamic Mechanical Analyzer (DMA) (Fig.63). The frequency dependant damping nature is evaluated.

Fig. 63: DMA testing and loss factor characteristics of CNT aligned nanocomposite

3B. Structural Testing and Evaluation (STE) Activities:

As a part of certification requirements, several structural testing and evaluation work was taken up by the group. Structural testing and evaluation of LCA mid board pylon fairing and engine mount structure of SARAS-PT1N aircraft were completed successfully. Further preparations for static structural strength testing of Rustum-II fuselage and damage tolerance testing of LCA wing root fitting box continued. Also, new proposals for structural testing of Rustum-II wing and empennage, life extension of MiG-29 landing gears were submitted to sponsors.

NEAT

EPOXYR-0.1% R-0.2% A-0.2%A-0.1% R-0.3% A-0.3%

Data Acquisition

Electronics

Electrodynamic

shaker

Beam

Structural testing of LCA- composite mid board pylon fairing (aft):

The Light Combat Aircraft-Tejas has presently eight metallic fairings in the inboard and mid board stations. As a part of the weight reduction program, it was proposed to convert these metallic fairing to composites. The mid-board pylon fairing (aft) which was metallic earlier has been designed with carbon sandwich construction, made of Carbon/Epoxy skin and Nomex flex core. The fairing has overall dimension of 774×162×310 mm and is U-shaped sandwich section with a total thickness of 8.7 mm as shown in Fig.64(a). It has to undergo structural qualification test before its fitment on the aircraft. The pylon fairing which is attached to the bottom of the LCA wing will experience the aerodynamic loads during flight.

Test rigs and loading fixtures were designed and developed for carrying out the tests. Fig. 64(b) shows the test set where the fairing was attached to the rig through a mounting plate and L-clamps. The rig consists of 2 consecutive portal frames inter connected by box section beam members and supported by angular gussets to which the pylon fairing was attached. The portal frames are welded to the base frames which are bolted to the reaction floor. Nylon pressure bag with compressed air was used for applying pressure which was controlled by a pressure regulator with a digital pressure gauge for monitoring the applied pressure. The pressure was applied on the inner surface of the fairing with the nylon bag confirms to the shape of the fairing contour.

(a) Schematic of fairing (b) Test set-up Fig. 64. A schematic of mid board pylon fairing (aft) and structural

test set up for testing of LCA fairing

The structural testing was carried out for two load cases (limit and ultimate). In the load case-1, where the pressure of 5 KPa was applied to the pylon fairing on one half part of the component, In the Load case-2, where the fairing was pressurized for 27 kPa and 41 KPa respectively for both limit and ultimate load cases. The test article was strain gauged at 8 critical locations using stacked rosette gauges (R). Five (5) Cable Displacement Sensors (CDS) were used for the measurement of deflection at the identified critical locations. In the load case-1, the maximum principal strains were less than100 με and the maximum deflection measured was 0.4 mm. But in the Load case-2, it was observed that the maximum principal strains were 329 με and 435 με respectively at the strain gauge R2 for both limit and ultimate cases. Also, the maximum deflections measured during the limit load and ultimate load cases were 1.42 mm and 2.12 mm respectively. The structural testing was successfully carried out for all the limit and ultimate load cases. It was concluded that the overall strains and deflections monitored during the test were well below the design allowable limits.

Materials Testing and Evaluation (MTE) Activities:

The mechanical characterization of aerospace materials, both metallic alloys and polymer composites, were carried out under various sponsored projects from ADA, HAL, DRDO etc. Fatigue tests on aluminum alloys were conducted to determine fatigue properties. Fibre Reinforced Polymer (FRP) matrix composite materials were tested to determine their various mechanical properties such as tension, compression, flexure and inter-laminar shear stress under RT and hot-wet conditions.

Static limit load test on engine mount structure: SARAS-PT1N aircraft:

The engine mount structure of the SARAS PT1-N has been redesigned to have a fail-safe design to cater uncontained rotor burst requirement, for which the engine mount design has been modified with an introduction of fail-safe truss assembly between yoke assembly and L and M frames of nacelle. The newly designed nacelle is part of the engine mount structure to fulfill the requirement of sealing. It was required to qualify the newly designed engine mount of SARAS PT1-N aircraft by carrying out a static load test for a critical design limit load case namely Propeller Malfunction to clear for flights including first flight of SARAS-PT1N aircraft. The work was carried out by MSD group and FSIG provided all the testing support for this work. Static load test rig on the stub wing engine mount assembly is shown in Fig. 2. Loads on the stub-wing were applied using lead shots bags. Engine mount loads at four mounting points were applied through mechanical turn-buckle arrangements. Also, loads on the L and M frames were applied by turn buckle and pulley arrangements. The test article was strain gauged at critical locations using linear gauges (72 nos.) and rosettes (23 Nos.). Data acquisition system (System 5000 from Micro- measurement group USA) was used for acquiring the strain data from all the gauge locations. Deflections during the tests were measured using two Cable Displacement Sensors (CDS) at identified locations. The test was carried out successfully and the strains monitored during the test were well below the allowable limits. The test was witnessed by CEMILAC certification agency

Carbon fiber composite IM7/ 8552:

As a part of design data generation and derivation of design allowable, prospective structural carbon fiber composite IM7/ 8552 to be used in LCA was tested to determine various mechanical properties. Over 20 different types of mechanical tests were conducted. Tests were performed in both RT and hot-wet conditions. A typical hot-wet test set-up is shown in Fig. 65. Prior to hot-wet testing, specimens were hygrothermally aged in an environmental chamber until moisture absorption saturation was attained. All the tests were performed following their respective ASTM standards. Over 500 number of tests were conducted, test data analyzed. Compilation of all the test results, derivation of design allowable and report preparations are underway.

Fig. 65: A photograph showing hot-wet

testing of composite specimen

.--------------------------------------------------------------------------------------------------------------------------

4. Infrastructure and facility Activity

IVHM Activity

A. Electrical power system and wiring health management Developed Health management algorithms for single and twin engine aircraft electrical system (ES) using deterministic (state machine) method, with a provision to (i) Assess each subsystem health(ii)Display effect of one subsystem fault on the other in the ES and (iii)Detect incipient faults and failure modes of the subsystem. The above mentioned algorithms with sensor data are implemented using LABVIEW in real time. A test rig for aircraft conversion system (TRU) is completed with sensor interface. In addition, bus power controller code of twin engine ES and simulation of 3 ø multilevel inverter for aircraft ES application are completed. Hardware with controller is in progress. A demonstration of single engine ESHM and twin engine ESHM along with aircraft TRU test rig and multilevel inverter with control signals is made ready utilizing the available HILS (PXI & opal RT) platforms. Located aircraft generators/IDGs and procured 30-40KVA IDGs (3 nrs) of Russian origin from ADA on permanent basis. The assembly, installation, commissioning and testing the units are being planned with the help of ADA & HAL teams. Also tested the TRU procured from HAL Lucknow, at ARDC for acceptance. Initiated work on Electrochemical impedance spectroscopy to find the health of Ni-CD battery.

Landing Gear Health Management Completed modeling and simulation of 30 nrs of Line Replaceable Units of the Landing gear System, of which two models are with fault injection methodology. These are hydraulic simulation with faults (i) with actuator internal leakage (ii) Check valve Final verification process is in progress. Developed a diagnostic algorithm related to the event of increase in the time of under carriage extension for landing gear health management using synthetic data. Developed 3D simulation models of extension and retraction of NLG & MLG with provision for fault injection with the following features: (i) NLG mechanism simulation with up-lock for retraction, (ii) NLG mechanism simulation with down-lock for retraction and (iii) NLG mechanism simulation with extension and retraction with hydraulics in co-simulation mode.

Prepared a scheme for condition monitoring of a hydraulic actuation system which can be used for nose landing wheel steering mechanism simulation in HILS platform. Got the integrated HILS and MBSE platform commissioned. Final approval is in progress.

--------------------------------------------------------------------------------------------------------------------------------------