Mach number is a dimensionless quantity in fluid dynamics representing the ratio of flow velocity past a boundary to the local speed of sound. It is named after the Moravian physicist and philosopher Ernst Mach.
A ramjet, or athodyd, is a form of airbreathing jet engine that uses the forward motion of the engine to produce thrust. Since it produces no thrust when stationary ramjet-powered vehicles require an assisted take-off like a rocket assist to accelerate it to a speed where it begins to produce thrust. Ramjets work most efficiently at supersonic speeds around Mach 3 and can operate up to speeds of Mach 6.
Wind tunnels are large tubes with air blowing through them which are used to replicate the interaction between air and an object flying through the air or moving along the ground. Researchers use wind tunnels to learn more about how an aircraft will fly. NASA uses wind tunnels to test scale models of aircraft and spacecraft. Some wind tunnels are large enough to contain full-size versions of vehicles. The wind tunnel moves air around an object, making it seem as if the object is flying.
A sonic boom is a sound associated with shock waves created when an object travels through the air faster than the speed of sound. Sonic booms generate enormous amounts of sound energy, sounding similar to an explosion or a thunderclap to the human ear. A decibel is the primary unit measurement of sound. "A thunderclap is incredibly loud, producing levels between 100 and 120 dBA - the equivalent of standing near a jet during take-off."
Compressible flow is the branch of fluid mechanics that deals with flows having significant changes in fluid density. While all flows are compressible, flows are usually treated as being incompressible when the Mach number is smaller than 0.3. The study of compressible flow is relevant to high-speed aircraft, jet engines, rocket motors, high-speed entry into a planetary atmosphere, gas pipelines, commercial applications such as abrasive blasting, and many other fields.
A scramjet is a variant of a ramjet airbreathing jet engine in which combustion takes place in supersonic airflow. As in ramjets, a scramjet relies on high vehicle speed to compress the incoming air forcefully before combustion, but whereas a ramjet decelerates the air to subsonic velocities before combustion using shock cones, a scramjet has no shock cone and slows the airflow using shockwaves produced by its ignition source in place of a shock cone. This allows the scramjet to operate efficiently at extremely high speeds.
In fluid dynamics, the lift coefficient is a dimensionless quantity that relates the lift generated by a lifting body to the fluid density around the body, the fluid velocity and an associated reference area. A lifting body is a foil or a complete foil-bearing body such as a fixed-wing aircraft. CL is a function of the angle of the body to the flow, its Reynolds number and its Mach number. The section lift coefficient cl refers to the dynamic lift characteristics of a two-dimensional foil section, with the reference area replaced by the foil chord.
In fluid dynamics, a Mach wave is a pressure wave traveling with the speed of sound caused by a slight change of pressure added to a compressible flow. These weak waves can combine in supersonic flow to become a shock wave if sufficient Mach waves are present at any location. Such a shock wave is called a Mach stem or Mach front. Thus, it is possible to have shockless compression or expansion in a supersonic flow by having the production of Mach waves sufficiently spaced. A Mach wave is the weak limit of an oblique shock wave where time averages of flow quantities don't change;. If the size of the object moving at the speed of sound is near 0, then this domain of influence of the wave is called a Mach cone.
The true airspeed of an aircraft is the speed of the aircraft relative to the air mass through which it is flying. The true airspeed is important information for accurate navigation of an aircraft. Traditionally it is measured using an analogue TAS indicator, but as the Global Positioning System has become available for civilian use, the importance of such air-measuring instruments has decreased. Since indicated, as opposed to true, airspeed is a better indicator of margin above the stall, true airspeed is not used for controlling the aircraft; for these purposes the indicated airspeed – IAS or KIAS – is used. However, since indicated airspeed only shows true speed through the air at standard sea level pressure and temperature, a TAS meter is necessary for navigation purposes at cruising altitude in less dense air. The IAS meter reads very nearly the TAS at lower altitude and at lower speed. On jet airliners the TAS meter is usually hidden at speeds below 200 knots (370 km/h). Neither provides for accurate speed over the ground, since surface winds or winds aloft are not taken into account.
In fluid dynamics, drag is a force acting opposite to the relative motion of any object moving with respect to a surrounding fluid. This can exist between two fluid layers or between a fluid and a solid surface. Unlike other resistive forces, such as dry friction, which are nearly independent of velocity, the drag force depends on velocity.
An oblique shock wave is a shock wave that, unlike a normal shock, is inclined with respect to the incident upstream flow direction. It will occur when a supersonic flow encounters a corner that effectively turns the flow into itself and compresses. The upstream streamlines are uniformly deflected after the shock wave. The most common way to produce an oblique shock wave is to place a wedge into supersonic, compressible flow. Similar to a normal shock wave, the oblique shock wave consists of a very thin region across which nearly discontinuous changes in the thermodynamic properties of a gas occur. While the upstream and downstream flow directions are unchanged across a normal shock, they are different for flow across an oblique shock wave.
Choked flow is a compressible flow effect. The parameter that becomes "choked" or "limited" is the fluid velocity.
A hypersonic wind tunnel is designed to generate a hypersonic flow field in the working section, thus simulating the typical flow features of this flow regime - including compression shocks and pronounced boundary layer effects, entropy layer and viscous interaction zones and most importantly high total temperatures of the flow. The speed of these tunnels vary from Mach 5 to 15. The power requirement of a wind tunnel increases linearly with its cross section and flow density, but cubically with the test velocity required. Hence installation of a continuous, closed circuit wind tunnel remains a costly affair. The first continuous Mach 7-10 wind tunnel with 1x1 m test section was planned at Kochel am See, Germany during WW II and finally put into operation as 'Tunnel A' in the late 1950s at AEDC Tullahoma, TN, USA for an installed power of 57 MW. In view of these high facility demands, also intermittently operated experimental facilities like blow-down wind tunnels are designed and installed to simulate the hypersonic flow. A hypersonic wind tunnel comprises in flow direction the main components: heater/cooler arrangements, dryer, convergent/divergent nozzle, test section, second throat and diffuser. A blow-down wind tunnel has a low vacuum reservoir at the back end, while a continuously operated, closed circuit wind tunnel has a high power compressor installation instead. Since the temperature drops with the expanding flow, the air inside the test section has the chance of becoming liquefied. For that reason, preheating is particularly critical.
A supersonic expansion fan, technically known as Prandtl–Meyer expansion fan, a two-dimensional simple wave, is a centered expansion process that occurs when a supersonic flow turns around a convex corner. The fan consists of an infinite number of Mach waves, diverging from a sharp corner. When a flow turns around a smooth and circular corner, these waves can be extended backwards to meet at a point.
A Machmeter is an aircraft pitot-static system flight instrument that shows the ratio of the true airspeed to the speed of sound, a dimensionless quantity called Mach number. This is shown on a Machmeter as a decimal fraction. An aircraft flying at the speed of sound is flying at a Mach number of one, expressed as Mach 1.
Fanno flow is the adiabatic flow through a constant area duct where the effect of friction is considered. Compressibility effects often come into consideration, although the Fanno flow model certainly also applies to incompressible flow. For this model, the duct area remains constant, the flow is assumed to be steady and one-dimensional, and no mass is added within the duct. The Fanno flow model is considered an irreversible process due to viscous effects. The viscous friction causes the flow properties to change along the duct. The frictional effect is modeled as a shear stress at the wall acting on the fluid with uniform properties over any cross section of the duct.
Shock is an abrupt discontinuity in the flow field and it occurs in flows when the local flow speed exceeds the local sound speed. More specifically, it is a flow whose Mach number exceeds 1.
The University of Texas at Arlington Aerodynamics Research Center (ARC) is a facility located in the southeast portion of the campus operated under the Department of Mechanical and Aerospace Engineering. It was established in 1986 as part of an expansion of UTA's College of Engineering. The ARC contributes to the vision of UTA and the University of Texas System to transform the university into a full-fledged research institution. It showcases the aerodynamics research activities at UTA and, in its history, has established itself as a unique facility at a university level. The wind tunnels and equipment in the facility were mainly built by scouting for and upgrading decommissioned equipment from the government and industry. Currently, Masters and Ph.D. students perform research in the fields of high-speed gas dynamics, propulsion, and Computational fluid dynamics among other projects related to aerodynamics.
In gas dynamics, the Kantrowitz limit refers to a theoretical concept describing choked flow at supersonic or near-supersonic velocities. When an initially subsonic fluid flow experiences a reduction in cross-section area, the flow speeds up in order to maintain the same mass-flow rate, per the continuity equation. If a near supersonic flow experiences an area contraction, the velocity of the flow will increase until it reaches the local speed of sound, and the flow will be choked. This is the principle behind the Kantrowitz limit: it is the maximum amount of contraction a flow can experience before the flow chokes, and the flow speed can no longer be increased above this limit, independent of changes in upstream or downstream pressure.
