NASA/TM—2010-216259 AIAA–2009–5369

Wind Tunnel Model Design for the Study of Plume

Effects on Sonic Boom for Isolated Exhaust Nozzles

Raymond S. Castner

Glenn Research Center, Cleveland, Ohio

April 2010
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated to the

advancement of aeronautics and space science. The

NASA Scientific and Technical Information (STI)

program plays a key part in helping NASA maintain

this important role.

The NASA STI Program operates under the auspices

of the Agency Chief Information Officer. It collects,

organizes, provides for archiving, and disseminates

NASA’s STI. The NASA STI program provides access

to the NASA Aeronautics and Space Database and

its public interface, the NASA Technical Reports

Server, thus providing one of the largest collections

of aeronautical and space science STI in the world.

Results are published in both non-NASA channels

and by NASA in the NASA STI Report Series, which

includes the following report types:

TECHNICAL PUBLICATION. Reports of

completed research or a major significant phase

of research that present the results of NASA

programs and include extensive data or theoretical

analysis. Includes compilations of significant

scientific and technical data and information

deemed to be of continuing reference value.

NASA counterpart of peer-reviewed formal

professional papers but has less stringent

limitations on manuscript length and extent of

graphic presentations.

• TECHNICAL MEMORANDUM. Scientific

and technical findings that are preliminary or

of specialized interest, e.g., quick release

reports, working papers, and bibliographies that

contain minimal annotation. Does not contain

extensive analysis.

• CONTRACTOR REPORT. Scientific and

technical findings by NASA-sponsored

contractors and grantees.

• CONFERENCE PUBLICATION. Collected

papers from scientific and technical

conferences, symposia, seminars, or other

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific,

technical, or historical information from

NASA programs, projects, and missions, often

concerned with subjects having substantial

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and

technical material pertinent to NASA’s mission.

Specialized services also include creating custom

thesauri, building customized databases, organizing

and publishing research results.

For more information about the NASA STI

program, see the following:

• Access the NASA STI program home page at

http://www.sti.nasa.gov

• E-mail your question via the Internet to help@

sti.nasa.gov

• Fax your question to the NASA STI Help Desk

at 443–757–5803

• Telephone the NASA STI Help Desk at

443–757–5802

• Write to:

NASA Center for AeroSpace Information (CASI)

7115 Standard Drive

Hanover, MD 21076–1320
NASA/TM—2010-216259 AIAA–2009–5369

Wind Tunnel Model Design for the Study of Plume

Effects on Sonic Boom for Isolated Exhaust Nozzles

Raymond S. Castner

Glenn Research Center, Cleveland, Ohio

Prepared for the

45th Joint Propulsion Conference and Exhibit

cosponsored by AIAA, ASME, SAE, and ASEE

Denver, Colorado, August 2–5, 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

April 2010
Acknowledgments

The author recognizes the contributions of Harold Kingsberg, from Case Western Reserve University, who

performed signiÞ cant analysis supporting this design and test. In addition, Martin Silverman, John Jones,

and Warren Holt provided design and fabrication of the wind tunnel model and support systems.

Facility support was provided by Scott Smrdel.

This work was sponsored by the Fundamental Aeronautics Program

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management.

Available from

NASA Center for Aerospace Information National Technical Information Service

7115 Standard Drive 5301 Shawnee Road

Hanover, MD 21076–1320 Alexandria, VA 22312

Available electronically at http://gltrs.grc.nasa.gov
Wind Tunnel Model Design for the Study of Plume Effects

on Sonic Boom for Isolated Exhaust Nozzles

Raymond S. Castner

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind

Tunnel (SWT), with a goal to reduce the disturbance caused by supersonic aircraft flight over populated areas. This

work focused on the shock wave structure caused by the exhaust nozzle plume. Analysis and design was performed

on a new rig to test exhaust nozzle plume effects on sonic boom signature. Test capability included a baseline nozzle

test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum. Analysis was performed

on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls

caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to

Mach 3.0 with nozzle pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia).

The model design was based on a 1 in. outer diameter, was 9 in. in overall length, and was mounted in the wind

tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the strut was built to supply 90 psia of

pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia. Instrumentation was

developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field

static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between

two transparent sidewalls in the 1 ×1 SWT for Schlieren photography and comparison to CFD analysis.

Nomenclature

NPR Nozzle Pressure Ratio = Pt/P∞

P Local Static Pressure, psia

Pt Total Pressure in Nozzle

P. Free stream Static Pressure

AP P - P∞

X axial distance from Jet Simulator nosecone, inches

D Test nozzle diameter, inches

I. Introduction

NASA has conducted research programs to reduce or eliminate the operational restrictions of supersonic aircraft

over populated areas. Restrictions are due to the disturbance caused by the sonic boom, caused by the coalescence of

shock waves formed off the aircraft. Recent work has been performed to reduce the magnitude of the N-wave

generated by airplane components with focus on shock waves caused by the exhaust nozzle plume.

In previous work by Putnam and Capone (Ref. 1), a variety of nozzles were tested from a fully expanded Mach

1.7 nozzle to a fully expanded Mach 2.9 nozzle. Their study was conducted in a wind tunnel, where near field

pressure measurements were made at one nozzle diameter away from the model. Wind tunnel conditions were Mach

2.2 at a simulated altitude of 50,000 ft. Based on this work, a baseline nozzle was selected for a two-dimensional

Computational Fluid Dynamics (CFD) study.

During the CFD analysis, it was determined that over expanded and under expanded operation of the nozzle had

an effect on the N-wave boom signature; and demonstrated the feasibility of reducing the magnitude of the sonic

boom N-wave by controlling the nozzle plume interaction with the nozzle boat tail shock structure, as reported by

Castner (Ref. 2). Results from this CFD study were used to develop a combination of CFD and wind tunnel test

capability to validate and predict results obtained for supersonic nozzle concepts. A low cost test capability was

developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT) (Ref. 3). Analysis and

NASA/TM—2010-216259
design was performed for a new rig to test the exhaust nozzle plume effects on sonic boom signature. Test capability

included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper

plenum. Analysis was performed on the external and internal aerodynamic configuration.

II. Baseline Nozzle

A baseline nozzle was developed for initial checkout testing and qualification of the new rig in the 1×1 SWT.

Figure 1 depicts the design of the Mach 2.0 baseline supersonic exhaust nozzle configuration. Previous

computational and wind tunnel tests were performed on this nozzle at flight speeds of Mach 2.2 and an altitude of

50,000 ft. Figure 2 shows the larger scale wind tunnel model used by Putnam and Capone, where an x/D of zero is

the front of the nose cone. Near field pressure measurements were made at one nozzle diameter above the test

nozzle. Data taken from Putnam’s “Nozzle 6” was digitized from the report and displayed in Figure 3. A CFD data

set was obtained to match the experimental test conditions.

Pressure signatures, plotted as a ratio of ΔP/P∞, were compared to the experimental data. The results from the

comparison are plotted in Figure 3.

Initial wind tunnel testing in the 1×1 SWT will collect near field pressure data for comparison to the baseline

data. Upon successful validation, the new rig will be used to study novel concepts for nozzle plume effects on sonic

boom.

Following baseline nozzle testing, nozzle configurations were planned based on the NASA Dryden NF-15B Lift

and Nozzle Change Effects on Tail Shock research airplane. A study was also planned for supersonic nozzle pairs,

with varying levels of nozzle spacing and stagger.

Figure 1.—Baseline Nozzle.

NASA/TM—2010-216259 2
Figure 2.—Exhaust Nozzle Simulator from NASA TN-D-5553.

Figure 3.—Near Field Pressure Profile at 1 nozzle Diameter from baseline Nozzle 6.

NASAfTM—2010-216259 3
Nozzle

III. Analysis

Model layout included a support strut, upper plenum, nosecone, instrumentation, and nozzle mounting station,

(Fig. 4). To determine the design of this wind tunnel model, a study was performed to determine model size,

including both length and diameter. The objective was to have a region of 5 to 10 diameters in length downstream of

the nozzle exit, uninterrupted by reflected shock waves from wind tunnel walls. Model diameters of 1, 1.5, and 2 in.

were studied with spreadsheet models. Shock waves and the Prantdl-Meyer expansions were estimated from a conic

nose. The reflection path from wind tunnel walls was calculated. Analysis showed that a smaller model diameter

provided a longer reflection path for reflected shock waves from the tunnel walls. The longer reflection path

provided a longer axial distance before the shock waves intersected the tunnel centerline. As a result, a small model

diameter of 1 in. was selected. It was determined through analysis that a 15 ° (half angle) nosecone and a 3 in. long

plenum with a test nozzle would fit inside the reflected shock wave path, and still provide nearly 9 diameters of

nozzle plume for study.

A second study was performed to determine an optimal nose cone. In literature, there was much data on lift and

drag for nose cones, for example Reference 4. However, there was little published data for the shock structure

propagating from these nose cones. Nose cones were studied to see if the shock strength or reflection path would

change. Nose cones included a cone, a parabola, a power series, and the Von Karman nose cone, (Table 1). The

length of the nose cone was also varied from a maximum of 5.7 in. to a minimum of 2.85 in. in length.

Figure 4.—Wind Tunnel Model Cut-Away.

NASA/TM—2010-216259 4
TABLE 1.—NOSECONE DESIGNS (REF. 6)

Nosecone . R,h inches Equation

Long cone, 5° 0.5 5.7 y 

 xR

L



2(

x

) K( x )

2

y  R

11L

2  K

ll

Long 3/4 parabola 0.5 5.7  

 

 

Where K = 0.75

y  R

(L )

n

Long 3/4 power series 0.5 5.7

Where n = 0.75

y  

2  (L  x)2

 R 

Long ojive 0.5 5.7

R

2

 L

2

Where  =

2R

y  

2  (L  x)2

 R 

Short ojive 0.5 2.85

R

2

 L

2

Where  =

2R

y  R



)n

Short 3/4 power 0.5 2.85  L

Where n = 0.75



2



L

 K(

x

)2

y  R 

 2  K

)

Short 3/4 parabola 0.5 2.85 

 

 

Where K = 0.75

R  

sin(2 )

2

Short Von Karman 0.5 2.85

y





Where 0 = arccos(l – 1

L

2x



NASA/TM—2010-216259
Figure 5 depicts the CFD solutions for the long nosecones, 5.7 in. in length. The baseline conic nosecone with a

5° half angle had open space of 8.7 nozzle diameters downstream of the nozzle exit before the shock reflection

intersected the nozzle plume. Three other long nosecones had similar results, with the best nosecone being a 5°

cone.

8.7-D

Figure 5(a).—5 deg cone: 8.7 diameter between nozzle exit and reflected

shock wave at tunnel centerline.

Figure 5(b).—3/4 Parabola Long: 8.5 diameter between nozzle exit

and reflected shock wave at tunnel centerline.

•_ --- --- - _ 

_rte 8.5-D -

Figure 5(c).—3/4 Power Long: 8.5 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

Figure 5(d).—Long Ojive: 8.1 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

NASA/TM—2010-216259 6
Figure 6 shows the CFD solutions for the short nosecones, 2.85 in. in length. The short ojive was compared to

the long ojive in Figure 4(d). The useful space downstream of the nozzle exit increased to 9.6 nozzle diameters.

Three other nosecones had similar results with 10 diameters of useful space to observe nozzle plume shock

structure.

Mach Number: 1.9^ 1.975 2.05 2.125 2.2 _

9.6 -D

ANJEL

Figure 6(a).—Ojive Short: 9.6 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

10.0 -D

Figure 6(b).—3/4 Power Short: 10.0 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

Figure 6(c).—3/4 Parabola: 10.0 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

Figure 6(d).—Von Karman Short: 9.9 diameter between nozzle exit and

reflected shock wave at tunnel centerline.

NASA/TM—2010-216259 7
The first shock wave from the nosecone determined the location of the shock reflection. Shorter nosecones

caused stronger shock reflections due to higher angles of flow turning, with a slightly shorter reflection path;

however, the shorter path was offset by the shorter nosecone length. The end result was that all the short nosecones

provided more space to view nozzle plume shock structure. The short Von Karman nose cone was selected because

it was a low drag nose with good flow qualities.

The internal design of the model was driven by air flow requirements needed for “Nozzle 6”. A scaled version of

“Nozzle 6” required 0.109 lbm/sec of airflow at airspeed of Mach 2.0 and simulated altitude conditions of 50,000 ft.

A narrow strut with sharp leading and trailing edges was desirable to minimize tunnel blockage and shock wave

reflections in the lateral direction. The length of the strut needed to carry the airflow was the driver for the length of

the upper air plenum (Fig. 4); which was desired to be near 3 in., to match the length of the upper air plenum. A

3/8 in. strut thickness was selected with a 3 in. length, plus leading and trailing edges.

IV. Design

This wind tunnel model was designed to operate across a variety of wind tunnel Mach numbers, nozzle pressure

ratios, and simulated altitudes. Design Mach numbers could vary from 1.4 to 3.0. NPR could vary from 6 to 12, and

intended altitudes were 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia).

The model design was based on the 1 in. outer diameter, as was shown in Figure 4. The final length from the

nosecone to the nozzle exit was 9 in. Both nosecone and nozzle were attached with fine pitched threaded

connections. The support strut was 3/8 in. in diameter and contained nine passages to supply the needed airflow. Air

supply lines were 1/4 in. diameter, and the passages supply 90 psia of pressure to achieve 20 psia at the nozzle inlet,

for test conditions at simulated 50,000 ft altitude. To accommodate future larger nozzles and testing at lower

altitudes, the maximum design pressure at the nozzle inlet is 52 psia. Choked orifice holes were needed at the top of

the strut supply lines to obtain pressure and flow upstream of choked test nozzles. Orifice holes were 0.210 in.

diameter. Air was supplied to the bottom of the strut through a lower plenum by pressurized air systems at the 1 × 1

SWT. High pressure air was provided by the facility 450 psig supply system. This system was designed to provide

NPR settings within 1 percent of the set point.

The location of the model in the 1×1 SWT was determined by a shock wave pattern in the tunnel. This shock

wave was generated by a small step in the tunnel wall located at the end of the interchangeable wind tunnel throat

blocks, which were used to set tunnel Mach number. This tunnel shock reflection was located so as not to interfere

with the nozzle plume, in a similar fashion to the nosecone shock, (Fig. 7).

Figure 7.—Wind Tunnel Model Cross Section as installed in the 1x1 SWT test section.

NASA/TM—2010-216259
Figure 8.—Single Nozzle and Nozzle Pair Installation.

The location of the strut and model was designed to move forward and aft in the wind tunnel test section. This

was accomplished with a slot, allowing 6 in. of overall travel. In this manner the rig location could be modified to

avoid unforeseen wind tunnel flow effects or shock wave reflections. The design also included capability for testing

of exhaust nozzle pairs, simulating multiengine aircraft configurations. Two strut assemblies could be installed at

two specified distance increments, (Fig. 8). The first increment was a spacing of one nozzle diameter, center to

center (touching), and the second increment was three nozzle diameters. Both struts were supplied by a common air

manifold to provide matched test conditions.

V. Instrumentation and Data Systems

Instrumentation was located internally to measure nozzle pressure ratio, and was limited to one total pressure

probe upstream of the test nozzle, located on centerline, and one static pressure tap on the lower wall. An external

static pressure probe was designed to survey near field static pressure profiles at one nozzle diameter above the rig

centerline, also in Figure 7. The probe had a travel of 10 in. to capture pressure profiles starting from the nozzle exit

and extending back to where the plume interacts with the tunnel wall shock reflection. Two static pressure probes

were built based on Pinckney (Ref. 5), both a conventional and a short design. The conventional design, Figure 9,

had limitations where the static pressure measurements are not in the same flow environment as the probe tip. The

short probe design allowed for a shorter probe with less sensitivity to angle of attack. These probes were mounted on

a 3/8 in. thick airfoil strut, allowing 5.625 in. of probe length between the static pressure holes and the airfoil

support. Actuation of the probe was provided by a sliding actuator with stepper motor control. The actuator was

mounted outside the tunnel, in a sealed enclosure, and the airfoil support extended through a slot in the tunnel wall.

This model was placed between two transparent sidewalls in the 1 × 1 SWT for Schlieren photography and

comparison to Mach number contours generated by CFD analysis. Placement of the model in the 1×1 SWT was

determined by collecting Schlieren images with an empty test section. Figure 10 shows a Schlieren image from an

empty test section. Two shock wave patterns were present in an empty test section. These shock waves were traced

back to a joint in the wind tunnel wall at the end of the “Mach block”, which provided an interchangeable

convergent-divergent contour to set wind tunnel Mach number. These images were analyzed to determine the shock

wave angle of 29.4°. If the model nosecone was placed at the intersection of the shock waves, then the reflected

shocks from the “Mach block” would not reflect back to the wind tunnel centerline until 21.29 in. downstream of the

nose cone. This provided adequate space to study the nozzle plume shock structure without interference from

reflected shock waves.

NASA/TM—2010-216259 9
Figure 9.—Conventional Static Pressure Probe (Ref. 5).

Figure 10.—Schlieren image from empty test section.

Concluding Remarks

Analysis was conducted to design new capability for the 1 ×1 SWT. This new capability will provide a test rig to

study exhaust nozzle plume effects on sonic boom. Critical capabilities of this rig include the ability to test

interchangeable nozzles with a 1 in. outer diameter. Nozzle pressure ratios range from 6 to 12, and Mach number

capability ranges from 1.3 to 3.0.

Analysis was performed to size the nozzle test rig and avoid reflected shock waves from wind tunnel walls.

Studies included alternative nosecone designs to determine their effect on the reflected shock waves. Schlieren

images from an empty test section were analyzed to determine proper placement in the wind tunnel.

Plans include collection of static pressure profiles at one diameter above the tunnel centerline and Schlieren

images from exhaust nozzle plume effects.

References

1. Putnam, L. and Capone, F.: Experimental Determination of Equivalent Solid Bodies to Represent Jets

Exhausting into a Mach 2.20 External Stream, NASA TN-D-5553.

2. Castner, R.S.: Analysis of Plume Effects on Sonic Boom Signature for Isolated Nozzle Configurations,

NASA/TM—2008-215414.

3. Soeder, R., Stark. D., Leone, J., and Henry, M.: NASA Glenn 1- by 1- Foot Supersonic Wind Tunnel User

Manual, NASA/TM—1999-208478.

4. Seiff, A. and Sandahl, C.: The Effect of Nose Shape on the Drag of Bodies of Revolution at Zero Angle of

Attack, NASA RM-A51J25.

5. Pinckney, S.Z.: A Short Static Pressure Probe Design for Supersonic Flow, NASA TN D-7978.

6. “Nose Cone Design,” Wikipedia, 2009.

NASA/TM—2010-216259 10
REPORT DOCUMENTATION PAGE Form Approved

OMB No. 0704-0188

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302.

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB

control number.

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS.

1. REPORT DATE (DD-MM-YYYY) 2. REPORT TYPE 3. DATES COVERED (From - To)

01-04-2010 Technical Memorandum

4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER

Wind Tunnel Model Design for the Study of Plume Effects on Sonic Boom for Isolated

Exhaust Nozzles 5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S) 5d. PROJECT NUMBER

Castner, Raymond, S.

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 984754.02.07.03.13.05

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION

National Aeronautics and Space Administration REPORT NUMBER

John H. Glenn Research Center at Lewis Field E-17047-1

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITOR'S

National Aeronautics and Space Administration ACRONYM(S)

Washington, DC 20546-0001 NASA

11. SPONSORING/MONITORING

REPORT NUMBER

NASA/TM-2010-216259

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited

Subject Categories: 01, 02, 07, and 09

Available electronically at http://gltrs.grc.nasa.gov

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

A low cost test capability was developed at the NASA Glenn Research Center 1- by 1-Foot Supersonic Wind Tunnel (SWT), with a goal to

reduce the disturbance caused by supersonic aircraft flight over populated areas. This work focused on the shock wave structure caused by

the exhaust nozzle plume. Analysis and design was performed on a new rig to test exhaust nozzle plume effects on sonic boom signature.

Test capability included a baseline nozzle test article and a wind tunnel model consisting of a strut, a nosecone and an upper plenum.

Analysis was performed on the external and internal aerodynamic configuration, including the shock reflections from the wind tunnel walls

caused by the presence of the model nosecone. This wind tunnel model was designed to operate from Mach 1.4 to Mach 3.0 with nozzle

pressure ratios from 6 to 12 and altitudes from 30,000 ft (4.36 psia) to 50,000 ft (1.68 psia). The model design was based on a 1 in. outer

diameter, was 9 in. in overall length, and was mounted in the wind tunnel on a 3/8 in. wide support strut. For test conditions at 50,000 ft the

strut was built to supply 90 psia of pressure, and to achieve 20 psia at the nozzle inlet with a maximum nozzle pressure of 52 psia.

Instrumentation was developed to measure nozzle pressure ratio, and an external static pressure probe was designed to survey near field

static pressure profiles at one nozzle diameter above the rig centerline. Model layout placed test nozzles between two transparent sidewalls

in the 1 × 1 SWT for Schlieren photography and comparison to CFD analysis.

15. SUBJECT TERMS

Wind tunnel models; Wind tunnel tests; Models; Plumes; Sonic boom; Exhaust nozzles

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF 18. NUMBER 19a. NAME OF RESPONSIBLE PERSON

ABSTRACT OF

PAGES

STI Help Desk (email:help@sti.nasa.gov)

a. REPORT b. ABSTRACT c. THIS 19b. TELEPHONE NUMBER (include area code)

U U PAGE UU 16 443-757-5802

U

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18