Aerodynamics of delta wings with flaps at hypersonic speeds

Control surface effectiveness is an important parameter for any aeroplane. For a hypersonic aircraft, though the power required to operate the flaps is determined by low speed flying conditions, it is imperative to know the effect of flaps at hypersonic speeds. Hence, studies have been done on this topic by aerodynamicists for over 40 years. In spite of this, only a limited data is available in the literature on this subject. This paper discusses the experimental study of the effect of sweep on the aerodynamic characteristics of thin slab delta wings with flaps at hypersonic speeds. For the purpose of this investigation, a novel special thin six-component balance, which has a thickness of 4mm and can be housed inside wings with 8mm thickness, has been designed. The wings had a sweep of 76degrees, 70degrees and 65degrees, t/c of 0.053 and flaps with 12% of wing area and 12% of wing chord. Testing were done at Mach 8.2, Re number of 2.13 x 10(6) (based on chord), from alpha = -12degrees to 12degrees and flap angle of 20degrees, 30degrees and 40degrees. Separation lengths, measured from Schlieren pictures, clearly show that there is 'no appreciable' effect of sweep on them. Also, using a simple local flow field calculation, the separation has been identified to be transitional in nature. These features of separation reflect in the force data. Because of the small separation length, the flaps (inspite of their small size) were very effective in generating additional C-N, C-M and C-l, which increased with increase in flap angle. In general, the C-N, C-M and X-CP were unaffected by sweep for symmetric flap deflection at positive incidences and asymmetric flap case, For symmetric flap case at negative incidences, only C-N was not influenced by the sweep but C-M decreased and X-CP moved upstream as the sweep is decreased, The wing with lower sweep produces higher CA and lower (L/D)(max) for both symmetric and asymmetric flaps. The rolling moment and adverse yaw increased with decrease in sweep for asymmetric flap deflection. Newtonian theory is shown to be incapable of predicting the effect of sweep on C-l, C-n and on the incremental values of C-N, C-M and C-A. In conclusion, it can be said that a small flap is generally adequate for hypersonic aeroplanes provided they operate at altitudes where transitional and turbulent separation can be expected to occur. This would make the flaps effective and thus enable ample control authority.

Common-mode voltage eliminated three-level inverter using a three-level flying-capacitor inverter and cascaded H-bridge

This paper proposes a new 3 level common mode voltage eliminated inverter using an inverter structure formed by cascading a H-Bridge with a three-level flying capacitor inverter. The three phase space vector polygon formed by this configuration and the polygon formed by the common-mode eliminated states have been discussed. The entire system is simulated in Simulink and the results are experimentally verified. This system has an advantage that if one of devices in the H-Bridge fails, the system can still be operated as a normal 3 level inverter mode at full power. This inverter has many advantages like use of single DC-supply, making it possible for a back to back grid-tied converter application, improved reliability etc.

A Three-Level Common-Mode Voltage Eliminated Inverter With Single DC Supply Using Flying Capacitor Inverter and Cascaded H-Bridge

A three-level common-mode voltage eliminated inverter with single dc supply using flying capacitor inverter and cascaded H-bridge has been proposed in this paper. The three phase space vector polygon formed by this configuration and the polygon formed by the common-mode eliminated states have been discussed. The entire system is simulated in Simulink and the results are experimentally verified. This system has an advantage that if one of devices in the H-bridge fails, the system can still be operated as a normal three-level inverter at full power. This inverter has many other advantages like use of single dc supply, making it possible for a back-to-back grid-tied converter application, improved reliability, etc.

Collision avoidance between UAV clusters using swarm intelligence techniques

In this article, several basic swarming laws for Unmanned Aerial Vehicles (UAVs) are developed for both two-dimensional (2D) plane and three-dimensional (3D) space. Effects of these basic laws on the group behaviour of swarms of UAVs are studied. It is shown that when cohesion rule is applied an equilibrium condition is reached in which all the UAVs settle at the same altitude on a circle of constant radius. It is also proved analytically that this equilibrium condition is stable for all values of velocity and acceleration. A decentralised autonomous decision-making approach that achieves collision avoidance without any central authority is also proposed in this article. Algorithms are developed with the help of these swarming laws for two types of collision avoidance, Group-wise and Individual, in 2D plane and 3D space. Effect of various parameters are studied on both types of collision avoidance schemes through extensive simulations.

Experimental investigation of heat transfer reduction using forward facing cavity for missile shaped bodies flying at hypersonic speed

Forward facing circular nose cavity of 6 mm diameter in the nose portion of a generic missile shaped bodies is proposed to reduce the stagnation zone heat transfer. About 25% reduction in stagnation zone heat transfer is measured using platinum thin film sensors at Mach 8 in the IISc hypersonic shock tunnel. The presence of nose cavity does not alter the fundamental aerodynamic coefficients of the slender body. The experimental results along with the numerically predicted results is also discussed in this paper.

A Reactive Inverse PN Algorithm for Collision Avoidance among Multiple Unmanned Aerial Vehicles

While performing a mission, multiple Unmanned Aerial Vehicles (UAVs) need to avoid each other to prevent collisions among them. In this paper, we design a collision avoidance algorithm to resolve the conflict among UAVs that are on a collision course while flying to heir respective destinations. The collision avoidance algorithm consist of each UAV that is on a collision course reactively executing a maneuver that will, as in `inverse' Proportional Navigation (PN), increase Line of Sight (LOS) rate between them, resulting in a `pulling out' of collision course. The algorithm is tested for high density traffic scenarios as well as for robustness in the presence of noise.

Convective heat-transfer rate distributions over a missile shaped body flying at hypersonic speeds

Experiments are carried out with air as the test gas to obtain the surface convective heating rate on a missile shaped body flying at hypersonic speeds. The effect of fins on the surface heating rates of missile frustum is also investigated. The tests are performed in a hypersonic shock tunnel at stagnation enthalpy of 2 MJ/kg and zero degree angle of attack. The experiments are conducted at flow Mach number of 5.75 and 8 with an effective test time of 1 ms. The measured stagnation-point heat-transfer data compares well with the theoretical value estimated using Fay and Riddell expression. The measured heat-transfer rate with fin configuration is slightly higher than that of model without fin. The normalized values of experimentally measured heat transfer rate and Stanton number compare well with the numerically estimated results. (C) 2009 Elsevier Inc. All rights reserved.

Nonlinear aeroelasticity of rotating and flapping wings - a review

The interaction between large deflections, rotation effects and unsteady aerodynamics makes the dynamic analysis of rotating and flapping wing a nonlinear aeroelastic problem. This problem is governed by nonlinear periodic partial differential equations whose solution is needed to calculate the response and loads acting on vehicles using rotary or flapping wings for lift generation. We look at three important problems in this paper. The first problem shows the effect of nonlinear phenomenon coming from piezoelectric actuators used for helicopter vibration control. The second problem looks at the propagation on material uncertainty on the nonlinear response, vibration and aeroelastic stability of a composite helicopter rotor. The third problem considers the use of piezoelectric actuators for generating large motions in a dragonfly inspired flapping wing. These problems provide interesting insights into nonlinear aeroelasticity and show the likelihood of surprising phenomenon which needs to be considered during the design of rotary and flapping wing vehicle

A Simple Five-Level Inverter Topology for Induction Motor Drive Using Conventional Two-Level Inverters and Flying Capacitor Technique

This paper proposes a new five-level inverter topology for open-end winding induction motor (IM) drive. The popular existing circuit configurations for five-level inverter include the NPC inverter and flying capacitor topologies. Compared to the NPC inverter, the proposed topology eliminates eighteen clamping diodes having different voltage ratings in the present circuit. Moreover it requires only one capacitor bank per phase, whereas flying capacitor schemes for five level topologies require six capacitor banks per phase. The proposed topology is realized by feeding the phase winding of an open-end induction motor with two-level inverters in series with flying capacitors. The flying capacitor voltages are balanced using the switching state redundancy for full modulation range. The proposed inverter scheme is capable of producing two-level to five-level pulse width modulated voltage across the phase winding depending on the modulation range. Additionally, in case of any switch failure in the flying capacitor connection, the proposed inverter topology can be operated as a three-level inverter for full modulation range. The proposed scheme is experimentally verified on a four pole, 5hp induction motor drive.

Search Strategies for Multiple UAV Search and Destroy Missions

Multiple UAVs are deployed to carry out a search and destroy mission in a bounded region. The UAVs have limited sensor range and can carry limited resources which reduce with use. The UAVs perform a search task to detect targets. When a target is detected which requires different type and quantities of resources to completely destroy, then a team of UAVs called as a coalition is formed to attack the target. The coalition members have to modify their route to attack the target, in the process, the search task is affected, as search and destroy tasks are coupled. The performance of the mission is a function of the search and the task allocation strategies. Therefore, for a given task allocation strategy, we need to devise search strategies that are efficient. In this paper, we propose three different search strategies namely; random search strategy, lanes based search strategy and grid based search strategy and analyze their performance through Monte-Carlo simulations. The results show that the grid based search strategy performs the best but with high information overhead.

Multiple UAV Coalitions for a Search and Prosecute Mission

Unmanned aerial vehicles (UAVs) have the potential to carry resources in support of search and prosecute operations. Often to completely prosecute a target, UAVs may have to simultaneously attack the target with various resources with different capacities. However, the UAVs are capable of carrying only limited resources in small quantities, hence, a group of UAVs (coalition) needs to be assigned that satisfies the target resource requirement. The assigned coalition must be such that it minimizes the target prosecution delay and the size of the coalition. The problem of forming coalitions is computationally intensive due to the combinatorial nature of the problem, but for real-time applications computationally cheap solutions are required. In this paper, we propose decentralized sub-optimal (polynomial time) and decentralized optimal coalition formation algorithms that generate coalitions for a single target with low computational complexity. We compare the performance of the proposed algorithms to that of a global optimal solution for which we need to solve a centralized combinatorial optimization problem. This problem is computationally intensive because the solution has to (a) provide a coalition for each target, (b) design a sequence in which targets need to be prosecuted, and (c) take into account reduction of UAV resources with usage. To solve this problem we use the Particle Swarm Optimization (PSO) technique. Through simulations, we study the performance of the proposed algorithms in terms of mission performance, complexity of the algorithms and the time taken to form the coalition. The simulation results show that the solution provided by the proposed algorithms is close to the global optimal solution and requires far less computational resources.

Injection cooling of blunt bodies flying at high Mach numbers

A study of transpiration cooling of blunt bodies such as a hemicylinder is made by solving Navier-Stokes equations. An upwind, implicit time-marching code is developed for this purpose. The study is conducted for both perfect-gas and real-gas (chemical equilibrium) flows. Investigations are carried out for a special wall condition that is referred to as no heat flow into the wall condition. The effects of air injection on wall temperature are analyzed. Analyses are carried out for Mach numbers ranging between 6-10 and Reynolds numbers ranging between 10(6)-10(7). Studies are made for spatially constant as well as spatially varying mass injection rate distributions, White cold air injection reduces the wall temperature substantially, transpiration cooling is relatively less effective when the gas is in chemical equilibrium.

Shock tunnel study of spiked aerodynamic bodies flying at hypersonic Mach numbers

A three-component accelerometer balance system is used to study the drag reduction effect of an aerodisc on large angle blunt cones flying at hypersonic Mach numbers. Measurements in a hypersonic shock tunnel at a freestream Mach number of 5.75 indicate more than 50% reduction in the drag coefficient for a 120degrees apex angle blunt cone with a forward facing aerospike having a flat faced aerodisc at moderate angles of attack. Enhancement of drag has been observed for higher angles of attack due to the impingement of the flow separation shock on the windward side of the cone. The flowfields around the large angle blunt cone with aerospike assembly flying at hypersonic Mach numbers are also simulated numerically using a commercial CFD code. The pressure and density levels on the model surface, which is under the aerodynamic shadow of the flat disc tipped spike, are found very low and a drag reduction of 64.34% has been deduced numerically.

Studies on the effects of test gas on the flow field around large angle blunt cone flying at hypersonic Mach number

The effect of the test gas on the flow field around a 120degrees apex angle blunt cone has been investigated in a shock tunnel at a nominal Mach number of 5.75. The shock standoff distance around the blunt cone was measured by an electrical discharge technique using both carbon dioxide and air as test gases. The forebody laminar convective heat transfer to the blunt cone was measured with platinum thin-film sensors in both air and carbon dioxide environments. An increase of 10 to 15% in the measured heat transfer values was observed with carbon dioxide as the test gas in comparison to air. The measured thickness of the shock layer along the stagnation streamline was 3.57 +/- 0.17 mm in air and 3.29 +/- 0.26 mm in carbon dioxide. The computed thickness of the shock layer for air and carbon dioxide were 3.98 mm and 3.02 mm, respectively. The observed increase in the measured heat transfer rates in carbon dioxide compared to air was due to the higher density ratio across the bow shock wave and the reduced shock layer thickness.