Source
of
Acquisition
NASA
Marshall
Space
Flight
Center
The Dynamics
of
Shock Dispersion and Interactions in Supersonic Freestreams with
Counterflowing Jets
Endwell
0.
Daso", Victor E. Pritchett', Ten-See Wangm
NASA Marshall Space Flight Center, MSFC,
AL
35812, USA
Dale
K.
Ota'
HyPerComp, Inc., Westlake Village, CA
91362,
USA
Isaiah M. Blankson*
NASA Glenn Research Center,
Cleveland, OH 44135, USA
and
Aaron H. Auslender@
NASA Langley Research Center
Hampton, VA 23681-2199, USA
Abstract
An
active flow control concept using counterflowing jets to significantly modify the external flowfields and
strongly weaken or disperse the shock-waves of supersonic and hypersonic vehicles to reduce the aerothermal loads
and wave drag was investigated. Experiments were conducted in a trisonic blow-down wind-tunnel, complemented
by pre-test computational fluid dynamics (CFD) analysis of a 2.6% scale model of Apollo capsule, with and without
counterflowing jets, in Mach 3.48 and 4.0 freestreams, to assess the potential aerothermal and aerodynamic benefits
of this concept. The model was instrumented with heat flux gauges, thermocouples and pressure taps, and employed
five counterflowing jet nozzles (three sonic and other
two
supersonic with design Mach numbers of 2.44 and 2.94)
and nozzle exit diameters ranging from 0.25 to 0.5 inch. Schlieren data show that at low jet flow rates of 0.05 and
0.1
lb,/sec, the interactions result in a long penetration mode (LPM) jet, while the short penetration mode (SPM) jet
is observed at flow rates greater than
0.
Ilb,/sec., consistent with the pre-test CFD predictions. For the LPM, the jet
appears to be nearly hlly-expanded, resulting in a very unsteady and oscillatory flow structure in which the bow
shock becomes highly dispersed such that it is no longer discernable. Higher speed camera Schlieren data reveal the
shock to be dispersed into striations of compression waves, which suddenly coalesce to a weaker bow shock with a
larger standoff distance as the flow rate reached a critical value. The pronounced shock dispersion could significantly
impact the aerodynamic performance (L/D) and heat flux reduction of spacecrafk in atmospheric entry and re-entry,
and could also attenuate the entropy layer in hypersonic blunt body flows. For heat transfer, the results show
significant reduction in heat flux, even giving negative heat flux for some of the SPM interactions, indicating that the
flow wetting the model is cooling, instead of heating the model, which could significantly impact the requirements
and design of thermal protection system. These findings strongly suggest that the application of counterflowing jets
as active flow control could have strong impact on supersonic and hypersonic vehicle design and performance.
Introduction
One of the technical challenges in space exploration and interplanetary missions is controlled entry and re-entry
into planetary and Earth atmospheres, which requires the dissipation of considerable kinetic energy as the spacecrafk
decelerates and penetrates the atmosphere. As such, effective heat load management of stagnation points and
acreage heating remain a technological challenge and pose significant risk, especially for human missions.
@
AST, Aerospace Flight Systems, Senior Member, AIAA, E-mail: Endwell.O.Daso@nasa.gov
X
AST, Aerospace Engineer, Member, AIAA
Technical Assistant, Senior Member, AIAA
+
Member
of
Technical Staff, Senior Member, AIAA
*
Senior Technologist, Associate Fellow, AIAA
@I
Assistant Head, Hypersonic Airbreathing Propulsion Branch
This paper is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
1
American Institute of Aeronautics and Astronautics
The efficiency and performance of an aerospace vehicle or spacecraft in atmospheric flight are dictated by the
physics of the flowfield about the vehicle. Flowfields of supersonic and hypersonic vehicles are characterized by
strong shock waves, contributing disproportionately to the vehicle drag and aerothermal loads, which translate into
poor aerodynamic performance (liwdrag) and stringent thermal protection system (TPS) requirements, and other
performance penalties including vehicle range, weight and payload.
In this work, we
start
out with an extensive review of previous work in counterflowing or opposing jets for
vehicle drag and heat flux reductions in supersonic and hypersonic flows. This is followed with the discussion of the
results of the experimental and computational analyses we performed to gain a better understanding of the flow
physics in order to advance the technology readiness level of the application of counterflowing jets as a viable active
flow technology, and demonstrate the effects and potential benefits of the concept.
Backcround
Since the early 1950s there has been continued interest in the application of “active flow control” concepts to
modify or change the external flowfields of transonic, supersonic and hypersonic vehicles and spacecraft in order to
reduce wave drag and aerothermal loads, and for spacecraft deceleration. Thus active flow concepts to weaken the
shock systems received considerable interest in the
1950~’~.
The work of Stadler and Inouye’ showed that opposing
jets at low “flow weights” or flow rates doubled the convective heat transfer to a hemispherical model while the heat
transfer was reduced by half with tangentially injected jets at equivalent flow weights at the stagnation point. On the
other hand, the results of Rashis3 showed considerable surface cooling with counterflowing water jets. Ferri and
Bloom4 conducted analysis, based on an approximate theory, and tests of upstream water and air injection in a Mach
6.1 fieestream with five model shapes, fiom a sphere-cylinder to cone-cylinders with slightly different heights. Their
study showed that directed upstream fluid injection could be effective means of cooling in hypersonic flow and the
theoretical analysis correlated well with the air jet cooling on the
50”
cone model. Resler and Sears’ explored the use
of electromagnetic effects to improve aerodynamic performance and Ziemer6 demonstrated the effects of magnetic
fields on the standoff distance of the bow shock of a sphere in a supersonic stream, at the same time showing the bow
shock to be significantly diffused, with an attendant increase in shock standoff distance.
Studies were also done on various flowfield modifications for wave drag and heat load reduction, and also as
potential improvements for radio wave attenuation as a solution for communications blackout, in the 1960s and
1970~~-’~, using both gas and liquid counterflow or forward-facing jets, including a flight test experiment”. Warren7
conducted an experimental study of the effect of ejecting nitrogen and helium gases upstream into a Mach
5.8
fieestream of a sphere-cone model. The coolant gases effectively reduced the heat flux on the model if the
fieestream was not disturbed by the injected gas, that is, at low coolant gas flow rates, while the coolant effectiveness
was considerably reduced at larger flow rates. He also noted that with the injection, a “stagnation circle” formed on
the model with an attendant increase in the heat flux, resulting in a net heat flux greater than the case without any
injection.
Charczenco and Hennessey* employed a retrorocket to produce a supersonic counterflowing jet to investigate
drag reduction on a sphere with a conical aftbody model. They observed flow instability about the nose, and the
drag was reduced below the case with the retrorocket off except at very large retrorocket thrust coefficients. Romeo
and Sterrett9”’ investigated the effect of a forward-facing jet on the bow shock of bht body in a Mach 6 fieestream.
Their investigation revealed two modes of shock displacement: one in which the blunt body bow shock grew in size
but retained its structure, while in the second mode, the shock standoff distance increased considerably with the
shock becoming less steady with increasing Mach number and high total pressure ratios. Grimaud, and McRee”
performed stagnation-point gas injection experiment on a hemispherical-cone in a hypersonic arc tunnel. At lower
coolant flow rates, the blunt body heating rates increased initially, but decreased with increasing flow rate, with up to
a
33%
reduction in heat transfer for different coolants at the maximum coolant injection rates. Beckwith and
Bushnell” reported the results of the flight test experiment at
150,000
ft. altitude and a velocity of 14,000
Ws,
using
intermittent nose and side port water injections to reduce aerodynamic heating
of
a sphere-cone with 9” half-angle
cone-cylinder flare and a spherical blunt nose
(RAM
B2). They reported a decrease in surface temperature ranging
fi-om 155°F to
408T,
corresponding to heat loads of 240 to 470 BWf? over a period of 90 seconds due to the water
injection. From their data and water evaporation theory, they calculated negative surface heat transfer, indicating the
cooling effect of the pulsed side port injection, which progressively became less negative with distance fiom the port.
Barbed3 tested the cooling effects of counterflowing stagnation point injection with nitrogen, helium and
hydrogen on a
90”
sector copper hemisphere in a Mach 6-8 tunnel with total enthalpies between
1500
to
5000
BWlb,. For nitrogen and helium injections, the average heat transfer decreased initially with increasing “relative
mass flow,” while it increased by
10%
above the no injection case due to an impinging or reattached “stagnation
2
American Institute of Aeronautics and Astronautics
ring” at a “relative mass flow” of about 8. For hydrogen injection, they reported a 65% increase was found as a
result of combustion. However, as the relative mass flow increased, the average heat transfer decreased
considerably, becoming as low
as
10% of the case with no injection at large relative mass flows for all test gases.
Finley14 analytically and experimentally investigated counterflowing jets in a Mach 2.5 freestream using
two jet
nozzles: a sonic and a Mach 2.6 nozzle, having different diameter ratios with respect to the model. His detailed
analysis showed that the effects of the jet Mach number and “flow force coefficient,” are critical in determining
whether the flow is steady or unsteady. Keyes and Heker” conducted tests with “retro-propulsion”
as
counterflowing jets for spacecraft deceleration in atmospheric decent and showed that for jets located at the
periphery, the aerodynamic drag generally increased with jet total pressure.
Bushnell and Huffman16 studied long penetration jet interactions and noted that the cooling effects of stagnation
point injection depended on the penetration distance of the water jet. Jarvinen and Adams” showed that for single
nozzle retrorocket engine on conical aeroshell planetary
entry
vehicles, the flowfield had
two
regimes of jet
penetration; long penetration at low “thrusting” coefficients and a short blunt jet penetration which terminated in a
terminal shock14 at large coefficients. They also observed that the transition from long penetration to short
penetration occurred at fixed jet exit pressure to freestream pressure ratio for all engine sizes tested. McGhee” also
investigated the effects of centrally located “retronozzle,” with an exit Mach number of
3.0,
on a 140” blunt cone in
Mach 3.0,4.5 and 6.0 freestreams at
O”,
2” and
5”
angles of attack. From the observations of the flow physics, three
regimes were identified in terms
of
the nozzle jet expansion: Regime 1, where the jet was over-expanded with the
local flow static pressure greater than the jet exit pressure; Regime 2, where the jet was fully-expanded with the local
flow static pressure approximately equal to the jet exit pressure, and Regime
3,
where the jet was under-expanded,
with the local flow static pressure less than the jet exit pressure. For Regimes
1
and 2, the flowfield was unsteady,
while it was steady in Regime
3
at all angles of attack. In the steady flow regime, locations of the jet (barrel) shock,
flow interface and the bow shock were found to be primarily a function of nozzle thrust coefficient. Grenich and
Woods’’ demonstrated the concept for heat load and drag reduction on ballutes in a Mach 20 freestream, and also
identified regions of steady and unsteady flow for low and high jet flow rates”, respectively. Their test results show
significant reduction in the drag coefficient on the ballute as a function of jet flow rate.
In recent years, there has been a strong interest in the application of weakly ionized nonequilibrium plasma
(WINP) counterflowing jets for the reduction
of
wave drag and heat flux of bodies in supersonic and hypersonic
flows. More recent work20-40 have revealed that WINP jets into high-speed flows produce various shock-attenuating
and anomalous effects. Malmuth et and Formin, et studied different jet penetrations in plasma jet
experiments with truncated cone-cylinder models in Mach 2, 2.5, and 4 supersonic fi-eestreams. These experiments
also revealed
two
modes
of
jet-penetration: short penetration mode (SPM) and long penetration mode (LPM),
consistent with the previous findings of Refs. 17 and 18, with the LPM giving greater reduction in wave drag and
larger shock stand-off distance.
Shang, et and Shang39 have investigated experimentally and computationally the aerodynamic effects of
various counterflowing jets, with and without plasma, to determine the amount of drag reduction in a hypersonic
flow over a sphere. The jet penetration was observed to have
two
stability modes: an unsteady oscillatory motion
under a “subcritical” state and a nearly steady “supercritical” state beyond the shock bifurcation point, depending on
the driving stagnation pressure and mass flow rate of the jet. The drag reduction was found to depend strongly on
the jet mass flow rate, in agreement with previous work, and had the same trend with and without the plasma.
However, the plasma jet gave about 10% more reduction
in drag3’ which was attributed to the deposited thermal
energy of the plasma. Josyula et a1.4’ investigated the applications of counterflowing jets for drag reduction in high
speed systems, and came to the conclusion that it is most suited for hypersonic blunt nosed bodies. Gilinsky, et a1:l
conducted experiments and CFD analysis to modify the shock wave of cylindrical and “butt-end” blunt bodies in
subsonic and supersonic flows using single and multiple needles, and opposing liquid jets. Their results showed a
considerable reduction in drag, with the drag coefficient decreasing with flow rate.
Daso, et a1.4’ obtained a CFD solution for the SPM jet giving better than
15%
reduction in drag, accounting for
the jet thrust, from a sonic counterflowing jet of a truncated cone-cylinder in a Mach 2 freestream, and also
developed an analytical approach to establish a sustained LPM jet.
showed the effect of
counterflowing water jets
in
a Mach 6 fi-eestream, which significantly modified the strong bow shock on sphere-
cylinder to a oblique shock. Hayashi, et a1.@ undertook both numerical and experimental study of opposing jets in a
supersonic flow over a sphere-cylinder for thermal protection. Similar to previous work, their study also showed
considerable reduction both in drag and heat transfer. They also identified both steady regime and unsteady flow
with shock oscillations, depending on the ratio of the stagnation pressure
of
the jet to that of the freestream.
Woods, et
3
American Institute of Aeronautics and Astronautics
The above review shows the extensive body of work on the use of counterflowing or opposing jets to modify
supersonic and hypersonic flows for shock attenuation, and the reduction of wave drag and heat transfer. However,
in nearly all the previous work, the model geometries used were typically sphere-cylinders, sphere-cones, cylinders,
(truncated) cone-cylinders or a simple aeroshell. Thus, there is a paucity of work with models that are more
representative of actual spacecraft configurations such as re-entry vehicles.
In
this work, we investigated
experimentally and numerically the flow environment of supersonic freestream-counterflowing jet interactions, and
the attendant aerodynamic and aerothermal effects of an actual spacecraft model, a 2.6% scale Apollo capsule. We
discuss below the results of the analyses of the flow structure of the Apollo capsule model to gain a better insight of
the flow physics to advance the technology readiness of active flow control as a viable technology for spacecraft, and
supersonic and hypersonic vehicles for mitigating aerothermal loads and improving lift to drag ratio for enhanced
aerodynamic performance. Though the review of previous work above included counterflowing plasma jets in order
to discuss a broader body of work, the work reported here does not include weakly ionized plasmas jets or any
plasma effects. Thus, only “cold” counterflowing jets were employed in this study.
Model Geometrv and Instrumentation DescriDtion
The test model is a 2.6% scale Apollo capsule of
4”
diameter, designed to accommodate five axisymmetric
interchangeable counterflowing jet nozzles, as shown in Fig. 1. The secondary air supply feedline to the nozzles is
housed in the model sting (Fig. IC). The feedline is a 51’8” OD high pressure tube welded to the backside of the
nozzle socket and regulated with a maximum of 900 psi valve. A Schlieren flow visualization system with low and
high speed cameras was also used to visualize the flowfield and capture the interactions and resulting shock structure
and dynamics. Pressure, heat flux and temperature data were acquired with a newly installed Labview data
acquisition system. The instrumentation consists of 56 static pressure ports and 15 Medtherm Schmidt-Boelter heat
flux transducers/thermocouples (model no. 8-5-0.5-36-SE-20486, range: O-SBtu/ft?sec, with calibration uncertainty
of
+3%
responsivity) on both the heatshield and the conical aftbody, which is truncated slightly to admit the specially
designed sting. The gauges were installed
in
a circular pattern or rows, with each row having three or more gauges,
Fig. le. The model and the nozzles are made with 174 PH HI050 stainless steel. The nozzles are trapped in a bore
in the base and are held in place by jack-on screws. The design allows the nozzles to be readily replaced or
interchanged with another from the heatshield side or face of the model without the need for disassembly or the
removal of the model from the sting during testing. Figure
2
shows the three sonic and two Mach 2.44 and 2.94
supersonic nozzles. The blank was used for the baseline geometry (no injection). The diameters of the nozzle exits
vary from 0.25” to 0.5” to assess the effective of nozzle geometry. The nozzle countours were designed using an in-
house nozzle design code, ADAPT45, which uses the method of characteristics with boundary layer displacement
thickness correction.
Facility Description
The test facility used for the experiments is the Marshall Space Flight Centers (MSFC) trisonic wind tunnel
(TWT). The test siction of the tunnei has a lPXI4” cross-section and
is3l”
long. It‘is an intermittent blow-down
tunnel, which operates by high-pressure air flowing from storage tanks to atmospheric or vacuum conditions. The
test section provides Mach numbers ranging from 0.2 to 4.96. Mach numbers between 0.2 and 1.3 are obtained by
using a variable diffuser, while the transonic Mach numbers of 0.95 to 1.3 are achieved through the use of plenum
suction and perforated walls. A solid wall supersonic test section provides the entire range
of
Mach numbers from
2.74 to 4.96, with one set of automatically actuated contour nozzle blocks. A hydraulically controlled pitch sector
located downstream of the test section provided the capability of testing at angles-of-attack from -10” to +lo”. For
the tests performed in this study, the freestream Mach numbers were 3.48 and 4.0 for both the baseline geometry and
with counterflowing jet nozzles. At Mach numbers higher than 4.0, the tunnel could not establish uniform flow due
to blockage. The tunnel conditions for the Mach 3.48 and 4.0 freestreams were total pressure of 44.92 psi and 54.85
psi, total temperature of 581.38”R, and 575.84”R, and unit Reynolds no. of 4.88X106 and 4.67X106, respectively.
Test Run Matrix
The
run
matrix included a wide range of
jet
parameters and two freestream supersonic Mach numbers of 3.48
and 4.0. There were five counterflowjet nozzles: three sonic and two supersonic, with varying nozzle exit diameters
of 0.25”, 0.375” and 0.5” to determine the effect of the nozzle Mach number and geometry. The nozzles were
run
at
design flow rates of 0.05,
0.10,
0.25, 0.35 and 0.501bm/sec, the corresponding to nozzle stagnation pressure and
temperatures are given in Table
1,
as well as at three jet angles of attack of +5”,
0”
and -9” to determine the effects of
flow rate and the angle of attack on the flow structure and interactions.
4
American Institute of Aeronautics and Astronautics
a) Illustrative sketch showing
b)
3-D
Solid Model
internal design
c) Model with heat
flux
(visible) and pressure tap instrumentation.
Figure
1.2.6%
Scale model of the Apollo capsule
Blank for Baseline
0.25”
Dest 0.375” De~t
0.5”
Dent 0.5” Dent
0.5”
Defit
Geometry Sonic Nozzle Sonic Nozzle Sonic Nozzle Mach
2.44
Nozzle Mach
2.94
Nozzle
Figure
2.
Sonic and Supersonic counterflowing jet nozzles, and blank for baseline geometry.
5
American Institute
of
Aeronautics and Astronautics