The hypersonic Ludwieg tube HLB [1] generates an intermittent Ma=6 air flow for Unit Reynolds numbers up to 20 million. The test section size allows to test models of about 0.2m length. Instrumentation includes Schlieren, infrared thermography, and pressure gauges.
The Ludwieg tube HLB is a long pipe with a converging - diverging nozzle, from which the flow exits into the test section, diffuser and dump tank. A fast-acting valve is placed just before the nozzle. When the valve opens, an expansion wave travels upstream into the storage or driver tube. The expansion wave reflects at the tube end and runs back until it again reaches the nozzle. This limits the period of time during which one can observe quasi steady flow conditions in the nozzle and the hypersonic test section. Due to its simple flow scheme the Ludwieg tube generates high speed flows with high flow quality without the need to install a total pressure control device and a large settling chamber as with conventional blow-down facilities. The electrical power requirement of this wind tunnel is remarkably low and such are the total operating costs.
The HLB is designed to achieve the hypersonic Mach number 6 in the test section of 0.5 m diameter. The maximum Reynolds number depends on the pressure of the storage tube which is limited to 30 bar. The run time with steady flow conditions in the test section is given by the tube length of 17 m. Valve operation is performed by a pneumatically driven piston which closes the throat 160 ms after opening. The storage tube is divided into a cold and a heated section, in order to avoid condensation and to allow heat transfer measurements. The 6 m³ vacuum tank down-stream of the diffuser en-ables the start up of the hypersonic flow and keeps the back pressure low enough to avoid flow breakdown during run time. The HLB is equipped with a high-pressure compressor, high pressure air storage, and a capable vacuum pump. These systems allow short cycle times of tunnel operation with about 10 runs every hour.
The diverging nozzle is carefully contoured. The opening half angle varies from initially 7.5° to 3° at the nozzle end. That is, some flow expansion is maintained into the test section in order to keep the nozzle flow insensitive to deviations of the design point. Comparisons of flow computations (figure above) and measurements in the test section confirm that the non- uniformity of the pitot pressure in the core flow is about than +/- 1.2% which corresponds to Mach number variations of +/- 0.6%. The test section itself is a tube with constant diameter. A cavity is attached to the bottom, which contains the two-component traversing mechanism. Optical access is achieved using windows of 0.26 m diameter on both sides and on the top of the test section.
The tunnel is equipped with an eight-channel acquisition system used for static or pitot pressure measurements at a rate of 10 kHz. A rake with 6 pitot probes is mounted on the traversing mechanism to allow flow field investigations within the test section. Instantaneous total temperatures of the flow entering the nozzle are also measured. Flow visualisation is possible with a Schlieren system. This method makes use of the fact that density gradients change the optical refractive index of the gas. The Schlieren system is installed using an unusual one-sided optical scheme, due to space limitations in the building. A Indigo infrared camera system with 320 x 256 resolution at 340 Hz is available for temperature measurements. Research is underway to develop infrared heat transfer measurements. In addition, the instationary onset of flow in the HLB and its impact on the flow around a generic testing model is numerically investigated in depth in a separate project.
A Mach 3 supersonic nozzle has been designed based on the infrastructure of the HLB [3]. By replacing the Mach 6 nozzle with a Mach 3 nozzle, together with inserting an additional nozzle and a settling chamber, the new tunnel succeeds to produce a Mach 3 flow. The new nozzle shares the storage tube, test section, diffuser and damp tank with the original HLB. The new tunnel inherits the steady stagnation storage tube flow conditions and the high operation efficiency from the HLB. Compared to the preferred hot run of HLB, the Mach 3 tunnel has no problem with condensation at an ambient stagnation temperature, further improving the tunnel efficiency.
The jet simulation facility resembles an Ariane 5 launcher. With this facility it is possible to investigate afterbody and propulsive jet flows.
Figure 1 shows a sketch of the jet simulation facility. The working principle is similar to the HLB wind tunnel itself. Outside of the HLB test section is the 32m long heated storage tube. The diameter of the storage tube is 18.88mm and it can be pressurized up to 140bar and heated up to 900K. The rocket model is placed along the centerline of the HLB test section. A tandem nozzle, consisting of the first nozzle, the settling chamber and the second nozzle, is integrated into the rocket model. The second nozzle is an axisymmetric Truncated Ideal Nozzle (TIC) with a mean exit Mach number, Mae = 2.5, and a de = 43mm nozzle exit diameter. The diameter of the settling chamber is dSC = 39mm. A system of three perforated plates is integrated in the settling chamber to improve uniformity of the flow upstream of the second nozzle.
For scaling the Ariane 5 launcher the nozzle to body diameter ratio is used. For the Ariane 5 the nozzle diameter is dAriane = 2.094m, and the body diameter is DAriane = 5.4m and hence, the ratio is dAriane/DAriane = 0.388. The diameter of the cylindrical body is D = 108mm while the model nozzle lip thickness is 0.5mm. Therefore the ratio has been calculated for the inner and outer nozzle diameter. With the inner nozzle diameter dinner = 42mm the ratio is d/D = 0.389 and with the outer nozzle diameter douter = 43mm the ratio is d/D = 0.398. Note that the external nozzle fairing length to body ratio L/D = 1.2 represents the Ariane 5 value as well.
After the start of the facility the flow detaches in the first nozzle and a shock system is generated. Because of this shock system the flow is decelerated to subsonic speed, at the entry of the settling chamber. In the settling chamber the flow uniformity is improved with perforated plates designed to reduce total pressure and to work as flow straighteners, see Fig. 1. In the second nozzle the flow is accelerated to Mae = 2.5 at the nozzle exit.
[1] Estorf, M.; Wolf, T. und Radespiel, R.:
Experimental and numerical investigations on the operation of the Hypersonic Ludwieg Tube Braunschweig
5th European Symposium on Aerothermodynamics for Space Vehicles, 2004
[2] Wolf, T.; Estorf, M.; Radespiel, R.:
Simulation of the Time-Dependent Flow Field in the Hypersonic Ludwieg Tube Braunschweig
4th Atmospheric Reentry Vehicles & Systems, 2005.
[3] Wu, J.; Radespiel, R.:
Tandem nozzle supersonic wind tunnel design
Int. J. of Engineering Systems Modelling and Simulation, 2013 Vol.5, No.1/2/3, pp.8-18
[4] Stephan, S., Radespiel, R. and Müller-Eigner, R.:PROPULSIVE JET SIMULATION IN A HYPERSONIC LUDWIEG TUNNELDeutscher Luft- und Raumfahrtkongress 2012, Berlin.
[5] Stephan, S., Radespiel, R. and Müller-Eigner, R. :Jet Simulation Facility using the Ludwieg Tube Principle Proceedings of the 5th European Conference for Aerospace Sciences (EUCASS), Munich, Germany, July 1-5 2013.