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In a typical Launch Vehicle (LV), dynamic responses due to various flight events are estimated through Coupled Load Analysis (CLA) where the launch vehicle is coupled with a spacecraft. A launch vehicle is subjected to various loads during its flight due to engine thrust depletion / shut-off, thrust oscillation, wind and gust, maneuvering loads. In aerospace industry a standard CLA is performed by generating the mathematical model of launch vehicle and coupling it with reduced mathematical model of satellite and applying the boundary conditions. A CLA is a time consuming process as several flight instances and load cases need to be considered along with generation of structural dynamic model at each time instants. For every new mission, the satellites are mission specific whereas the launch vehicle and the loads remain unchanged. To take advantage of this fact, a new method called “Fast CLA through Reanalysis technique” is proposed in the present paper.

Abstract : In any human space flight program, safety of the crew is of utmost priority. In case of exigency during atmospheric flight, the crew is safely and quickly rescued from the launch vehicle using Crew escape system. Crew escape system is a crucial part of the Human Space flight vehicle which carries the crew module away from the ascending launch vehicle by firing its rocket motors (Pitch Motor (PM), Low altitude Escape Motor (LEM) and High altitude Escape Motor (HEM)). The structural loads experienced by the crew escape system during the mission abort are severe as the propulsive forces, aerodynamic forces and inertial forces on the vehicle are significantly high. Since the mission abort can occur at anytime during the ascent phase of the launch vehicle, trajectory profiles are generated for abort at every one second interval of ascent flight time considering several combinations of dispersions on various propulsive parameters of abort motors and aero parameters.

Heavy Commercial Road Vehicles (HCRVs) may be more susceptible to rollover incidents due to their higher centre of gravity position than passenger vehicles, and rollover is one of the significant causes of HCRV accidents. Therefore, variation in vehicle roll behaviour becomes crucial to the safety of an HCRV. Toe misalignment is a commonly observed phenomenon in HCRVs, and studying its impact on roll behaviour is important. In this study, the impact of the symmetric toe and thrust misalignment on the roll behaviour of an HCRV is analysed using IPG TruckMaker®, a vehicle dynamics simulation software. A ramp steer manoeuvre was used for the simulations, and the toe misalignment on a wheel was chosen from the range [-0.21°, 0.21°]. Variation in roll behaviour was quantified using the steering wheel angle at which one-wheel lift-off (OWL) occurred (SWAL).

The requirement of the current scenario is to identify the sustainable material and process it into acceptable properties for current applications. The natural fiber is a prime sustainable material having the properties of biodegradability, plenty of availability, economical and adequate physical-mechanical property. Sesbania rostrata fiber is extracted from the stem of Sesbania rostrata plant which is cultivated along with Turmeric plants on 1000 acres annually as a nitrogen fixation plant. The fiber-reinforced composite is a tailor made material by altering the fiber and polymer weight proportion to achieve desired properties for applications. The natural fiber is a promising material to replace synthetic fiber to transform the composite into biodegradable. The making of holes in the biocomposite by the secondary process is essential for the assembly operation.

Thrust measurement systems come in many sizes and shapes, with varying degrees of complexity, accuracy and cost . For the purposes of this information report, the discussions of thrust measurement will be limited to axial thrust in single-axis test systems.

In this research, the performance of two commercially available icephobic coatings is evaluated on an 81% scaled-down version of the Bell Flight APT 70 drone propeller. Tests are performed in an icing wind tunnel (IWT) under selected severe icing conditions to test the ice protection capability of coatings against both glaze and rime ice. Two different coating formulations are used, one is a polydimethylsiloxane (PDMS) acetoxy terminated coating, the other an epoxy-silicone. The coatings were briefly characterized in terms of their surface roughness, water contact angle and ice adhesion reduction factor compared to aluminum using the centrifugal adhesion test (CAT). Blade sets were prepared for both coatings and a third uncoated set was tested for reference purposes. Tests in the IWT were performed to simulate a true airspeed of 35 m/s and a constant propeller rotational speed of 5 500 RPM.

The reactionless drive is an internal momentum engine which until recently has been deemed impossible under the laws of physics. In this paper, the authors will extend the equation for reaction less propulsion = F=−μq2/6πcmr2v×dBdt+B×dvdt and derive an additional equation, which we call “The Sektet Equation” governing the system of motion, FSek=−μq2/6πcmr2∗2B∗dBdt. The results of the paper show that significant thrusts can be generated on relatively low voltages and energy inputs. It applies this equation to explain how NASA’s EM drive likely produces thrust via the “Sektet Equation” using a three circuit analysis of the Sektet Force.

This method covers electric outboards that are rated in terms of static thrust.

Abstract The electric propulsion system plays an important role during the operation of a satellite, i.e., maintaining the position of the north-south poles, adjusting the attitude, and transferring the orbit, where vector adjustment device is a key part of the system. We developed a new large-angle device to transfer thruster orbital, which has three driving motors and the failure of a single motor cannot affect the operation. The posture angle and linear pair displacement of this mechanism are simulated using forward and inverse kinematics solutions. In the following, the actual adjustment angle was measured with a three-coordinates measuring instrument and a gradiometer to compare with the simulated values. This design has been successfully applied in China’s asteroid exploration mission.

This paper will briefly review the idea of a reactionless drive. It will then analytically derive a reactionless drive based on a specific application of the Abraham-Lorentz force. Simulated data on such a drive and its key characteristics will be discussed. A benchtop working model of the theoretical drive derived in this paper will be presented. Most schemes for a reactionless drive being proposed today rely on emissions of ions, particles, or light; however while these schemes work in theory, they produce minuscule amounts of thrust, even at perfect conditions. No reactionless drive has ever been demonstrated to work in practice at any scale. When such a drive is developed, space travel will open up tremendously due to the fact that spacecraft will not have to carry propellant fuel while cruising. Findings demonstrate how this reactionless drive is mathematically feasible.

This paper describes the implementation, integration, testing and performance evaluation of compact and battery-less alternator with external regulator for diesel engine for avionics application. The key responsibility of this alternator is to generate 2.8kW power with 28V regulated power supply for various loads. The alternator has been integrated and installed on the diesel engine and further tested on dynamometer and thrust cradle with propeller combination. The alternator when used in conjunction with ACU (Alternator Control Unit) that is designed to boot strap field voltage during low speed operation, has the ability to self-excite. The alternator / ACU system has the ability to generate power even in the absence of battery voltage i.e. in battery less systems or those in which the battery is not always connected to the alternator. External voltage regulator has been used which minimizes ripple up to 1.0V. The alternator rpm ranges from 3000 to 10000 for generating maximum power.

Abstract Passengers would always like to reach their destinations with minimum commute time. Generating a higher thrust is a necessity. This implies that the turbomachinery associated with the power plant has to rotate faster and with higher efficiencies. However, high rotational speeds, mainly in the transonic regime, often lead to boundary layer separation, shocks, compressor stall, and surge. The current investigation is an attempt to reduce the abovementioned phenomena. It involves the performance study of a smoothened controlled diffusion airfoil (CDA) blade that has been optimized by “Multi-Objective Genetic Algorithm” (MOGA) by altering maximum camber location and stagger angle. Inlet pressure is varied from 15 kPa to 30 kPa and the angle of attack ranging from 40.4° to 56.4°. C48-S16-BS1 is validated and considered as the baseline profile, and all other blades are collated to this.

Abstract Jet engine hot parts (e.g., jet nozzle) are a crucial source of aircraft’s infrared (IR) signature from the rearview, in 1.9-2.9 μm and 3-5 μm bands. The exhaust nozzle design used in a jet aircraft affects its performance and IR signature (which is also affected just by performance) from the engine layout. For supersonic aircraft (typically for M ∞ > 1.5), a converging-diverging (C-D) nozzle is preferred over a convergent nozzle for optimum performance. The diverging section of the C-D nozzle has a full range of visibility from the rearview; hence, it was not considered a prudent choice for low IR observability. This theoretical study compares the IR signature of the C-D nozzle with that of the convergent nozzle from the rearview in 1.9-2.9 μm and 3-5 μm bands for the same thrust.

Reduction gears are very commonly used in the automotive and aviation industries. A propeller’s efficiency decreases rapidly as the speed of the blade tips nears the speed of sound. An engine reduction gear enables the engine to develop more torque while reducing the propeller’s revolutions per minute (RPM). This prevents the propeller’s efficiency from decreasing. This work deals with a detailed methodology for the design and analysis of a single-stage reduction gear. Custom pinion and wheel were modelled along with the engine assembly. A custom mount was designed and fabricated to allow for engine-reduction gear system integration. In this work, a 7.5 cubic centimeter (cc) single-cylinder glow engine is used. Bending and contact stress analysis was performed, and the results were compared with the calculated stress values. The torque demands were analyzed for propellers of different sizes at varying aircraft speeds.