NASA Scientific and Technical Aerospace Reports

schemes and schemes with physical time subiterations are subject to time-step limitations in practice that are removed by
pseudo time sub-iterations. Computations for the circular-arc airfoil indicate that a one-equation turbulence model predicts the
unsteady separated flow better than an algebraic turbulence model; also, the hysteresis with Mach number of the self-excited
unsteadiness due to shock and boundary-layer separation is well predicted.
Author
Accuracy; Iteration; Navier-Stokes Equation; Turbulent Flow; Unsteady Flow; Factorization
20040111298 North Carolina State Univ., Raleigh, NC, USA
Hypervelocity Aeroheating Measurements in Wake of Mars Mission Entry Vehicle
Hollis, Brian R.; Perkins, John N.; [1995]; 12 pp.; In English; 26th AIAA Fluid Dynamics Conference, 19-22 Jun. 1995, San
Diego, CA, USA
Contract(s)/Grant(s): NAGW-1331; NAG1-1663
Report No.(s): AIAA Paper 95-2314; Copyright; Avail: CASI; A03, Hardcopy
Detailed measurements of aerodynamic heating rates in the wake of a Mars-Pathfinder configuration model have been
made. Heating data were obtained in a conventional wind tunnel, the NASA LaRC 31’ Mach 10 Air Tunnel, and in a
high-enthalpy impulse facility, the NASA HYPULSE expansion tube, in which air and CO2 were employed as test gases. The
enthalpy levels were 0.7 MJ/kg in the Mach 10 Tunnel, 12 MJ/kg at Mach 9.8 for HYPULSE CO2 tests and 14 MJ/kg at Mach
7.9 for HYPULSE air tests. Wake heating rates were also measured on three similar parametric configurations, and forebody
heating measurements were made in order to facilitate CFD comparisons. The ratio of peak wake heating to forebody
stagnation point heating in the Mach 10 Tunnel varied from 7% to 15% depending on the freestream Reynolds number. In
HYPULSE, the ratio was ~5% for both air and CO 2. It was observed that an increase in the ratio of forebody corner radius
to nose radius resulted in a decrease in peak wake heating, and moved the peak closer to the base of the forebody. The wake
flow establishment process in HYPULSE was studied, and a method was developed to determine when the wake has become
fully established.
Author
Aerodynamic Heating; Computational Fluid Dynamics; Mars Missions; Wakes; Wind Tunnel Tests; Wind Tunnels;
Hypervelocity Flow
20040111299 NASA Langley Research Center, Hampton, VA, USA
Calculated Effects of Nitric Oxide Flow Contamination on Scramjet Performance
Fischer, Karen E.; Rock, Kenneth E.; [1995]; 13 pp.; In English
Report No.(s): AIAA Paper 95-2524; Copyright; Avail: CASI; A03, Hardcopy
The level of nitric oxide contamination in the test gas of the NASA Langley Research Center Arc-Heated Scramjet Test
Facility and the effect of the contamination on scramjet test engine performance were investigated analytically. The study was
conducted for standard facility conditions corresponding to Mach 6, 7, and 8 flight simulations. The analytically determined
levels of nitric oxide produced in the facility are compared with experimentally measured levels. Results of the analysis
indicate that nitric oxide levels range from one to three mole percent, which corroborates the measured levels. A three-stream
combustor code with finite rate chemistry was used to investigate how nitric oxide affects scramjet performance in terms of
combustor pressure rise, heat release, and thrust performance. Results indicate minimal effects on engine performance for the
test conditions of this investigation.
Author
Contamination; Nitric Oxide; Supersonic Combustion Ramjet Engines; Test Facilities; Aircraft Performance; Gas Flow
20040111300 NASA Langley Research Center, Hampton, VA, USA
Fluidic Control of Nozzle Flow: Some Performance Measurements
Federspiel, John; Bangert, Linda; Wing, David; Hawkes, Tim; [1995]; 8 pp.; In English; Copyright; Avail: CASI; A02,
Hardcopy
Results are presented of an experimental program that investigated the use of a secondary air stream to control the amount
of flow through a convergent-divergent nozzle. These static tests utilized high pressure, ambient temperature air that was
injected at the throat of the nozzle through an annular slot. Multiple injection slot sizes and injection angles were tested. The
introduction of secondary flow was made in an opposing direction to the primary flow and the resulting flow field caused the
primary stream to react as though the physical throat size had been reduced. The percentage reduction in primary flow rate
was generally about twice the injected flow rate. The most effective throttling was achieved by injecting through the smallest
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