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The characteristics and behavior of a 0.3m transonic cryogenic wind tunnel are discussed. The wide band of usable Reynolds numbers is analyzed along with a flow visualization technique using propane. The combination of magnetic suspension with the cryogenic wind tunnel is described. An outline of the circuit showing the locations of the magnet system and the features of the tunnel are presented.
Inadequate Reynolds number similarity in testing of scaled models affects the quality of aerodynamic data from wind tunnels. This is due to scale effects of boundary-layer shock wave interaction which is likely to be severe at transonic speeds. The idea of operation of wind tunnels using test gas cooled to cryogenic temperatures has yielded a quantrum jump in the ability to realize full scale Reynolds number flow similarity in small transonic tunnels. In such tunnels, the basic flow control problem consists of obtaining and maintaining the desired test section flow parameters. Mach number, Reynolds number, and dynamic pressure are the three flow parameters that are usually required to be kept constant during the period of model aerodynamic data acquisition. The series of activity involved in modeling, control law development, mechanization of the control laws on a microcomputer, and the performance of a globally stable automatic control system for the 0.3-m Transonic Cryogenic Tunnel (TCT) are discussed. A lumped multi-variable nonlinear dynamic model of the cryogenic tunnel, generation of a set of linear control laws for small perturbation, and nonlinear control strategy for large set point changes including tunnel trajectory control are described. The details of mechanization of the control laws on a 16 bit microcomputer system, the software features, operator interface, the display and safety are discussed. The controller is shown to provide globally stable and reliable temperature control to + or - 0.2 K, pressure to + or - 0.07 psi and Mach number to + or - 0.002 of the set point value. This performance is obtained both during large set point commands as for a tunnel cooldown, and during aerodynamic data acquisition with intrusive activity like geometrical changes in the test section such as angle of attack changes, drag rake movements, wall adaptation and sidewall boundary-layer removal. Feasibility of the use of an automatic Reynolds number control mode with fixed Mach number control is demonstrated.
A flow visualization technique for the large wind tunnels of the National Full Scale Aerodynamics Complex (NFAC) is described. The technique uses a laser sheet generated by the NFAC Long Range Laser Velocimeter (LRLV) to illuminate a smoke-like tracer in the flow. The LRLV optical system is modified slightly, and a scanned mirror is added to generate the sheet. These modifications are described, in addition to the results of an initial performance test conducted in the 80- by 120-Foot Wind Tunnel. During this test, flow visualization was performed in the wake region behind a truck as part of a vehicle drag reduction study. The problems encountered during the test are discussed, in addition to the recommended improvements needed to enhance the performance of the technique for future applications.
Wind-tunnel tests were conducted on a baseline scramjet inlet model having fixed geometry and swept leading edges at M = 2.30, 2.96, 3.95, and 4.60 in the Langley unitary plan wind tunnel. The unit Reynolds number of the tests was held constant at 6.56 million per meter (2 million per foot). The objectives of the tests were to establish inlet performance and starting characteristics in the lower Mach number range of operation (less than M = 5). Surface pressures obtained on the inlet components are presented, along with the results of the internal flow surveys made at the throat and capture stations of the inlet. Contour plots of the inlet-flow-field parameters such as Mach numbers, pressure recovery, flow capture, local static and total pressure ratios at the survey stations are shown for the test Mach numbers.
The feasibility of simulating propulsion-induced aerodynamic effects on scaled aircraft models in wind tunnels employing Magnetic Suspension and Balance Systems. The investigation concerned itself with techniques of generating exhaust jets of appropriate characteristics. The objectives were to: (1) define thrust and mass flow requirements of jets; (2) evaluate techniques for generating propulsive gas within volume limitations imposed by magnetically-suspended models; (3) conduct simple diagnostic experiments for techniques involving new concepts; and (4) recommend experiments for demonstration of propulsion simulation techniques. Various techniques of generating exhaust jets of appropriate characteristics were evaluated on scaled aircraft models in wind tunnels with MSBS. Four concepts of remotely-operated propulsion simulators were examined. Three conceptual designs involving innovative adaptation of convenient technologies (compressed gas cylinders, liquid, and solid propellants) were developed. The fourth innovative concept, namely, the laser-assisted thruster, which can potentially simulate both inlet and exhaust flows, was found to require very high power levels for small thrust levels.
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Tests at Mach 6 and 8 have been conducted on a model of a wind tunnel sound radiation shield which consists of interchangeable circular and wedge-shaped longitudinal rods with gaps between the rods for boundary layer removal. In view of the large effective sweep angle of the circular rods with respect to the local flow, the boundary layer behavior was surprisingly well represented by swept cylinder theory. Laminar flow was maintained up to a maximum local unit Reynolds number of 8 million per foot which would presumably be independent of model length for a given rod diameter by analogy with results on swept cylinders. Reduction in wind-tunnel sound levels within the shielded region of the circular rod model were about 90% of the maximum possible reduction for an ideal shield.
- Problem Statement: - Magnetic Suspension Balance System (MSBS) allows operation of wind tunnels without sting effect - Small-scale MSBS system has been demonstrated; NASA is building a full-scale MSBS wind tunnel for supersonic speeds - MSBS testing can provide data that addresses mission needs; current data reconstruction methods do not combine all measurement data - Reconstruction methods for aerodynamic coefficients and their uncertainties are piecemeal, and in their infancy - Mission Need: - Enhancement of Entry, Descent, and Landing (EDL) technology during atmospheric flight is a top priority - Accurate dynamic aerodynamic coefficients are necessary for stability and control of entry vehicles - Goal: - Develop an Extended Kalman Filter (EKF) based framework to do reduction of MSBS data that will estimate aerodynamics, trajectory, and uncertainty information, combining data sources from the tunnel
A design program is developed which incorporates a modern approach to the design of supersonic/hypersonic wind-tunnel nozzles. The approach is obtained by the coupling of computational fluid dynamics (CFD) with design optimization. The program can be used to design a 2D or axisymmetric, supersonic or hypersonic, wind-tunnel nozzles that can be modeled with a calorically perfect gas. The nozzle design is obtained by solving a nonlinear least-squares optimization problem (LSOP). The LSOP is solved using an iterative procedure which requires intermediate flowfield solutions. The nozzle flowfield is simulated by solving the Navier-Stokes equations for the subsonic and transonic flow regions and the parabolized Navier-Stokes equations for the supersonic flow regions. The advantages of this method are that the design is based on the solution of the viscous equations eliminating the need to make separate corrections to a design contour, and the flexibility of applying the procedure to different types of nozzle design problems.
The rotor studied in the wind tunnel had previously been flight tested on a BO-105 helicopter. The investigation was conducted to determine the rotor's aeroelastic stability characteristics in hover and at airspeeds up to 143 knots. These characteristics are compared with those obtained from whirl-tower and flight tests and predictions from a digital computer simulation. It was found that the rotor was stable for all conditions tested. At constant tip speed, shaft angle, and airspeed, stability increases with blade collective pitch setting. No significant change in system damping occurred that was attributable to frequency coalescence between the rotor inplane regressing mode and the support modes. Stability levels determined in the wind tunnel were of the same magnitude and yielded the same trends as data obtained from whirl-tower and flight tests.
A full-scale winglet flight test on a KC-135 airplane with an upper winglet was conducted. Data were taken at Mach numbers from 0.70 to 0.82 at altitudes from 34,000 feet to 39,000 feet at stabilized flight conditions for wing/winglet configurations of basic wing tip, 15/-4 deg, 15/-2 deg, and 0/-4 deg winglet cant/incidence. An analysis of selected pressure distribution and data showed that with the basic wing tip, the flight and wind tunnel wing pressure distribution data showed good agreement. With winglets installed, the effects on the wing pressure distribution were mainly near the tip. Also, the flight and wind tunnel winglet pressure distributions had some significant differences primarily due to the oilcanning in flight. However, in general, the agreement was good. For the winglet cant and incidence configuration presented, the incidence had the largest effect on the winglet pressure distributions. The incremental flight wing deflection data showed that the semispan wind tunnel model did a reasonable job of simulating the aeroelastic effects at the wing tip. The flight loads data showed good agreement with predictions at the design point and also substantiated the predicted structural penalty (load increase) of the 15 deg cant/-2 deg incidence winglet configuration.
This report gives an overview of a recent improvement project to reduce the background noise in the NASA Glenn Research Center 9-by 15-Foot Low Speed Wind Tunnel. Noise surveys of the original tunnel configuration indicated that flow surface noise in the test section and noise propagating from other tunnel regions were the primary background noise contributors. Several novel features were incorporated in the design of the upgraded tunnel, including a low noise flow surface and serpentine baffles for noise reduction, along with acoustically treated turning vanes and a new diffuser. Three companion papers present implementation details including the acoustic design and the aerodynamic quality/design CFD. The upgrades reduced background noise in the test section by about 10dBA over a wide range of flow speeds and frequencies. It is hoped this will be a useful summary to aid future users of the facility and wind tunnel engineers everywhere.
A comprehensive study of braze alloy selection process and strength characterization with application to wind tunnel models is presented. The applications for this study include the installation of stainless steel pressure tubing in model airfoil sections make of 18 Ni 200 grade maraging steel and the joining of wing structural components by brazing. Acceptable braze alloys for these applications are identified along with process, thermal braze cycle data, and thermal management procedures. Shear specimens are used to evaluate comparative shear strength properties for the various alloys at both room and cryogenic (-300 F) temperatures and include the effects of electroless nickel plating. Nickel plating was found to significantly enhance both the wetability and strength properties for the various braze alloys studied. The data are provided for use in selecting braze alloys for use with 18 Ni grade 200 steel in the design of wind tunnel models to be tested in an ambient or cryogenic environment.
This paper summarizes the thermal analysis activity being performed in support of the design of wind tunnel models for the National Transonic Facility (NTF) at the NASA Langley Research Center. The goal of the analysis effort has been to address model design difficulties associated with the severe thermal environment of this cryogenic wind tunnel. The unique characteristics of this environment are discussed for various phases of tunnel operation. The methods used to calculate temperatures and thermal stresses in the models are also described. The results indicate that thermal considerations do not drive the design to the extent originally envisioned. In general, the problems identified are localized and associated with thermal mismatch between components. Several specific examples of thermal design problems and proposed solutions are given.
Low speed wind tunnel performance tests of two advanced propellers were completed. The 62.2 cm diameter adjustable pitch models were tested at Mach numbers typical of takeoff, initial climbout, and landing speeds in the 10 by 10 ft Supersonic Wind Tunnel. Both models had eight blades and a cruise design point operating condition of 0.80 Mach number, 10.668 km S.A. altitude, 243.8 m/s tip speed and a high power loading of 301 kW sq m. No adverse or unusual low speed operating conditions were found during the test with either the straight blade SR-2 or the 45 deg swept SR-3 propellers. The 45 deg swept propeller efficiency exceeded the straight blade efficiency by 4 to 5%. Typical net efficiencies of the straight and 45 deg swept propeller at a Mach 0.20 takeoff condition were 50.2 and 54.9% respectively. At a Mach 0.34 climb condition, the efficiencies were 53.7 and 59.1%. Reverse thrust data indicates that these propellers are capable of producing more reverse thrust at Mach 0.20 than a high bypass turbofan engine at Mach 0.20.
Low speed wind tunnel performance tests of two advanced propellers were completed. The 62.2 cm diameter adjustable pitch models were tested at Mach numbers typical of takeoff, initial climbout, and landing speeds in the 10 by 10 ft Supersonic Wind Tunnel. Both models had eight blades and a cruise design point operating condition of 0.80 Mach number, 10.668 km S.A. altitude, 243.8 m/s tip speed and a high power loading of 301 kW sq m. No adverse or unusual low speed operating conditions were found during the test with either the straight blade SR-2 or the 45 deg swept SR-3 propellers. The 45 deg swept propeller efficiency exceeded the straight blade efficiency by 4 to 5 percent. Typical net efficiencies of the straight and 45 deg swept propeller at a Mach 0.20 takeoff condition were 50.2 and 54.9 percent respectively. At a Mach 0.34 climb condition, the efficiencies were 53.7 and 59.1 percent. Reverse thrust data indicates that these propellers are capable of producing more reverse thrust at Mach 0.20 than a high bypass turbofan engine at Mach 0.20.
The feasibility of a Mach 10 to 20, high pressure electrothermal wind tunnel was assessed. A heater based on a continuous high power electric arc discharge capable of heating air to temperatures above 10,000 K and pressures of 15,000 atm is the key element of this wind tunnel. Results of analytical study indicate that the facility is capable of simulation conditions suitable for hypervelocity airbreathing propulsion testing up to Mach 16. In this case simulation was limited by pressure containment, high nozzle throat heat flux rates, and chemical freezing in the nozzle. The high total pressure capability improved the recombination chemistry in the facility nozzle as chemical equilibrium prevailed to the freezing point. Steady arc discharges were observed with liquid nitrogen flowing into the arc chamber during tests based on the two millisecond test facility. The measured steady pressure in the arc chamber was 4559 psi, which is two times greater than maximum total pressure obtainable in conventional arc heaters.
The characterstics of NITRONIC 40 were investigated in connection with its use in cryogenic wind tunnel models. In particular, the effects of carbide and sigma-phase precipitation resulting from heat treatment and the presence of delta ferrite were evaluated in relation to their effects on mechanical properties and the potential consequences of such degradation. Methods were examined for desensitizing the material and for possible removal of delta ferrite as a means of restoring the material to its advertised properties. It was found that heat treatment followed by cryogenic quenching is a technique capable of desensitizing NITRONIC 40. However, it was concluded that it is extremely difficult, if not impossible, to remove the delta ferrite from the existing stock of material. Furthermore, heat treatments for removing delta ferrite have to take place at temperatures that cause very large grain growth. The implications of using the degraded NITRONIC 40 material for cryogenic model testing were reviewed, and recommendations were submitted with regard to the acceptability of the material. The experience gained from the study of NITRONIC 40 clearly identifies the need to implement a policy for purchasing top-quality materials for cryogenic wind tunnel model applications.