APPENDIX A
- 11 - APPENDIX A ICING Tln�NEL FACILITIES The amount of time required for both aircraft and equipment main tenance to obtain sufficient data in flight toward the solution of icing problems, together with the difficulty of obtaining data in specified, controlled icing conditions, resulted in the design and construction of an icing wind tunnel at the NACA Lewis laboratory in 1943-1944.
The Lewis icing tunnel is a single-return closed-throat tunnel, the general arrangement of which is shown in figures 12 and 26. The tunnel is constructed of steel plate and is insulated with a 3-inch thickness of Fiberglas. The outer nonstructural shell covering the insulation is made of 1/8-inch steel sheets.
The tunnel is anchored at each end of the test section and at each end of the drive motor and supported by columns and sliding expansion joints at all other points in order to allow movement due to temperature stresses. The over-all size of the tur..nel shell is about 198 feet long and about 75 feet wide. The test section and a portion of the entrance cone and diffuser are surrounded by a steel housing to provide space for the essential test equipment and operating perscnnel. This space is called the test chamber. Because the test section is vented to the chamber, the air pressure decreases in the chamber during operation of the tunnel (normally .less thfm 3 ir.,. Hg). Ao. a:�::- 1::-,::k is provided to permit access by personnel to the chamber during a rur..
The test chamber contains three floor levels: the 7c;.1nd floor c.m:.
) taining electrical, thermocouple, water, a...�d balance-scale equipment; the second floor, containing the test section of the tu:-.,nel and the various controls, manometers, recording i�str��ents, and associated equipme�t; Personnel and the third floor, containing auxiliary measuring equipment.
access to the tu...-ir�el is generally from the second floor of the test chamber, while models are lowered into the test section from the third floor through a removable 48- by 140-inch access hatch in the roof of the test section.
The test section is rectangular in shape, 9 feet wide, 5 feet high, and 20 feet long. The air enters the test section from a large rect�gu lar section giving a contraction ratio of abc·.it 14 to 1. The test secticn of the tunnel is provided with a turntable on which models can be mounted, as well as side-wall trunnion mour.its. The maximum tunnel airspeed with icing conditions and a large model in the test section is 260 knots.
Windows are provided on both si�es a..�d in the rocf of the tUJL.�el The windows test section to allow observation of models durir.g a test.
in the tunnel sides are laminated, electrically heated units similar to - 12 - windshields o n many aircraft, while the win.dews in the test secti o n r oo :r are unheated. The p o wer supplied t o new wind o ws currently being installed is 500 watts per square :r oo t. The temperature o :r the plastic inner layer o :r the wind o w is detected by a nickel wire element. This element is used in a bridge circui� t o c o ntr o l the wind o w temperature.
Turning vanes are used in all right-angle c o rners o :r the tunnel. The vanes d o wnstream o :r the test section and ahead o :r the drive :ran are steam heated t o prevent icing.
The drive m o tor f o r the tunnel devel o ps 4160 h o rsep o wer. The drive c o nsists o :r a d o ubly :red wound-r o t o r inducti o n m o t o r. P o wer is supplied directly t o the stat o r, while the power f o r the r o t o r is supplied by a f o ur-machine variable and fixed frequency setup. A variable-speed d-c m o t o r, driven acc o rding t o the Ward-Le o nard system, drives an a-c genera t o r. The generat o r supplies the power t o the tunnel drive-m o t o r r o t o r.
The speed o f the drive m o t o r is g o verned by the speed o :r the a-c genera t o r o r d-c m o t o r, the speed o :r these machines being c o ntr o lled by varying the v o ltage t o the d-c m o t o r. The drive m o t o r has a speed range from 0 t o 540 rpn. A 200,000-cubic-:r oo t-per-minute, SO-h o rsep o wer bl o wer is used t o c oo l the drive m o t o r.
The tunnel drive m o t o r is c o upled directly t o a 25-:r oo t-diameter drive :ran with 12 blades. The :ran blades are w oo den, with the leading edges o f the blades pr o tected by ne o prene abrasi o n sh o es. Stati o nary c o ntra-vanes are used ahead o :r the :ran.
A ventilating �ewer is located downstream o :r the drive m o t o r. This t o wer permits an exchange o :r tunnel air with o utside air. The primary use o f this unit has been t o provide an addi tio."lSl. c o oling l o ad t o help regulate the tunnel air temperature :r o r certain test c o nditi o ns. In ad diti o n, a finned-tube heat exchanger with a capacity o f 5,000,000 Btu per hour is available t o aid in regulating tb.e tunnel air temperature.
A compressi o n-type refrigerati o n system l o cated in a nearby building is used t o c oo l the tunnel air t o the required icing c o nditi o n. The tunnel is c oo led by passing the air o ver a bank o f refrigerated :f'imied heat ex changers l o cated in an area between the drive m o t o r and the tunnel spray system. The t o tal refrigerati o n capacity is ab o ut 7700 t o ns. The n o rmal c oo ling load f o r the icing tunnel requires from 1200 t o 2100 tons; however, this requirement varies with climatic c o nditi o ns. Air temperatures as ° l o w as -40 C can be o btained, alth o ugh most tests are c o nducted in the o ° range o f .5 t o -20 C.
Icing c o nditi o ns similar t o those enc o untered in the atm o sphere are created by a battery o :r air-water atomizing n o zzles. A view o f the spray system l oo king d o wnstream int o the test section is shown in figure 13.
The spray n o zzles are m o unted in six h o riz o ntal spray bars and l o cated t o - 15 - give a UJ?.iform cloud approximately 4- by 4-feet in the test section.
Cc.ltrols for the spray system are located in airfoil-shaped enclosures at one end of each strut.
A sketch of the air-water atomizing nozzles used in the spray system is shown in figure 27. The nozzle assembly consists of •air and water supply lines, steam line to prevent icing of the entire strut, and the spray nozzle (Inconel). Approximately 80 nozzles are used to obtain an adequate cloud in the tunnel. The nos.zles were specially developed for th� tunnel t� yield droplet sizes r�ng from a mean ef�ective size of 4 microns at low water flows to abou.t 20 microns at maximum water flows.
Air pressures of 60 to 80 pounds per square inch are used normally to atomize the water, while the water pressures range from a few pounds �bove t.."le. air-pressure values up to 140 po-:J.nds per square inch.
The large water-pressure values correspond to large water flows and large droplet sizes. The droplets produced by these nozzles are not unifrom in size but va:ry approximately in accordance with a Iangm.uir Dor E drop-pize distribution (ref. 57). For a constant water flow, ._the liquid water content in the tunnel varies with the airspeed. !n addition, the contraction of the tunnel entrance cone affects the droplet paths and local water concentration in the tunnel. Consequently, maximum or mini- .mum vaiues of water content and droplet size independent of tunnel air speed cannot be stated explicitly.
The water used for the spray system pa��es through a 500-gallon-per hour-capacity demineralizer. The demineralizer consists of two anion and cation filter beds tbat remove all :minerals from t..he water, thereby pre venting fouling and plugging of the spray system. From the demineralizer, the water is piped into a storage ta!lk with a capacity of 750 gallons.
The storage tank is kept fall by a floe.t- switch that turns the demineraliz er on and off to maintain a given wate:r level in the st..orage tank. Fram the storage tank the water is pip€?d to two turbine water pumps with a capacity of 5 gallons per minute ea.en. at l.50 pounds per square inch gage.
The water is pumped through three rota.m�ters for flow measurement and then ° into steam heat exchanger t!la.t heats the water to a temperature of 80 � to 90 C. Heating of the water is uecessary to prevent freeze-out of the water when it is air-atomized to cloud. d:rop.lets in the tunnel. Following the heat exchanger the water is filtered at each str11t control box. The water pressure is regulated at each str-�t by a pressure regulator con- .
trolled by the tunnel operator. !'he·water pressure is sensed by a pres sure transmitter, which changes water pressm:·e to pneumatic pressure. The water pressures in each spray strut {int.he form of J.)Il.eumatic pressure) are indicated on a manometer board in the test chamber.
Air is :f'urni.shed to the water-spray system f'ran a service air line with a capacity of 6 pou:ids per second at 120 pounds per square inch • This air is passed through a pressure regulator, a steam heat exchanger 0 °
(which heats the air to approximately so to 90 C), and a two-stage
filter befo�e entering the strut control boxes. The air pressure is also controlled from the test-chamber control area.
- 14 - A separate spray system consisving Of 4 to 9 nozzles is used to in ject dyed water .:l.nto tl1e tuc,>:_e,. for exp�rimer_:.a:� 3t:,:d.ies of droplet impi".!ge ment characteristics of various bo,iies (ref. 15) o A balance frame is provided w:ith a 6-component force-measuring scale system. Data are recorded automatically on tapes at each balance.
Electrically heated co-axial pressure tul:)es are used to obtain pressure data. All pressure data are recorded phc�ographically from multi-tube manometer boards. Temperature data obtained with copper-constantan thermo The control equipment couples are recorded on automatic flight recorders, includes variable transformers for power control to models, automatic tem perature controller for heated air to models, and various recording instru ments for heat-source control. Standard icstruments are used to record tunnel airspeed and ai= temperature. Ar. NACA pressure-type icing-rate meter is used to measure the liquid-water content of the twmel atmosphere.
Special instrumentation is added whenever required for a particular study.
Heated air for providing models with icing protection is supplied by three heat exchangers. The air from these exchangers is heated by the ex haust from a Jet-engine combustion can. Each heat exchanger has a flow capacity of 1000 pounds per hour with pressure regulation up to 120 pounds per square inch.
Orifices in each line allow a measurement of the flow Constant air temperature over a wide range of air from each exchanger.
flows is obtained by an automatic flow control that regulates the amount of cold air penni tted to mix with heated air from the e�c:':1.angers.
Electric heating supplied to models for icing protection can be ob tained from either a-c or d-c sources; however; a-c is generally preferred.
The d-c system capacity is 28 volts and rated a� 100 am�eres. In addition, a 12-volt d-c system rated at 50 B.!!1.peres is also available. A 29-volt a-c system rated at 50 amperes is ava:!.lable and is used for heater studies for which the heater load is normal}.�,r rtm on d.co A 110-volt single-phase system and a 208-volt, three-phase 50-ampere system are available for J large electrical loads. Selective power inputs (a-c three-phase system) to electrically heated models are metered (power recorded on a recording wattmeter) by means of 18 variable transfc-r.mers rated at 3 amperes, 16 variable transformers rated at 9 amperes, a.�d 3 variable transformers 0 1 rated at 45 amperes. A 400 cycle :5..: ve:-t-:::r.- capable of s ;.pplying 1500 volt "'.
amperes at 115 volts is also used for some instrument tests.
Electronic timers are ava:.la"ble by which specifiei heating and icing periods for either electric or air heating systems can be controlled.
The 136-inch-high multi-tube manometer board is so arranged that it may operate as an integrating type er a star:d.ard board. A total of 298 readings can be obtained from the ooard. Additional U-tube and standard manometer boards are available as required. Most of the tubes are also connected to an air-purge system by which air is bled through the tubes - 15 - back to the model. An air-operated cylinder or pincher closes off the tubes at the manometer to prevent the purge air from blowing the manom eter fluid out of the bonds. The purge air prevents the entry of water from the spray cloud into the tubes and blocking or freezing of the un heated portions of the l'l"essure·lir-es. During this purging procedure, no manometer-board readings are taken.
For aerodynamic studies of airfoils in icing conditions, the airfoil surfaces aft of the region protected by the icing protection system (called afterbody) are generally heated. This heating is required, since the turbulence level and. supersaturated air in the test section cause a frost deposit on the cold portions of a model. T'�ese deposits have rarely been observed in natural flight icing. Such frost deposits increase the measured model drag. A steam line operating at +5 to -5 inches of mercury is used to heat these afterbodies. To avoid steam leakage from the mod.el into the tunnel, the afterbody is operated at a negative pressure �Y means of a small ejector and a barometric condenser located externally of the model.
Photographs of ice formations during a test are obtained with high speed electronic flash equipment, while conventional camera equipment is used for pictures taken in the tunnel at the conclusion of a test. Color photography has proved to be the most satisfactory for movie film record ing of data in the presence of the spray cloud.
Airfoil models nonnally span the vertical height of the tunnel.
Chords of these models have ranged f'ran 15 to 96 inches or larger.
Horizontal model mounting has also been used occasionally; however, be cause of wind-tunnel-wall interference effects, the vertical mounting is preferred. Bodies of revolution and inlets tes�ed. are no:nr.ally less than 56 inches in diameter.
In addition to the 6- by 9-foot icing tunnel, two s,maller high-speed icing-duct tunnel £acilities are al.so used. Techniques equivalent to those just described are used in these tunnels. A schematic diagram of the 5.84- by 10-inch tunnel presented in figure 28 shows the inlet dif fuser section with screens, the plenum er.amber with flow-straightening tubes, the bellmoutl:. tunnel entry, t:ie test section, ai:.d the outlet dif fuser section. The tunnel is designed to provide a range of subsonic Ma.ch numbers from 0.5 to 0.8 and. a supersonic Mach number of 2.0. Altitudes up to 30,000 feet may be simulated.
° A suppiy of refrigerated air initially at approximately -20 F and.
-4 with a specific humidity of s.muo pound of water:�"!!' pound of dry air is conditioned to provide the desired temperatures and humidities at the tw:mel test section.
The humidity of the airstream is controlled by means of steam injected at a point sufficiently far upstream to ensure thorough mixing at the tunnel entry.
- 16 - One wall of the.tunnel contains a large glass section for observa tion and visual measurements. The other wall has five porthole.s for access to·the inside of the tunnel and removable plugs for installation 9f instrumentation at various stations along the tunnel� ·. Permanent in strumentation' of the tunnel at the teat section includes static-pressure -:ta,ps along the topr and bottom surfaces of the tunnel and pressure taps and·�ermocouplea ::tri the plenum chamber •
. . . '
· ''.A• schematic diagram of the canponent parts of a 1� ... by 12-inch icing :duct tunnel is shown in figure 29. The tunnel characteristics are in general ·similar to the 3.84- by 10-inch tunnel. Subsonic ,speeds up to. a
Mach ��oe� of 0.75 can be achieved with.small airfoil mode�s." _For icing
studie•• the airspeed is maintained in the subsonic-and 10'«,ijuperaonic speed �e�.
.. .
-.17 - APPE.?-.1DIX B OPERATION� TECHNIQUES The ·following are techniques used. for model testing 1n'the Lewis .
icing. tunnel and associated facilities.
Thermocouple Installations Whenever possible, all skin or surface thermocouples (copper constantan) are peened into small holes drilled into the surface as shown 111 figure 50. The ball at the junction of the thermocouple is just large enough to fit 'into the hole, so �bat peening the surface around the ball
will· resuJt in a firmly anchored thermocouple. The ball should be as
. clos·e to ·the outer ·surface of the skin as possible. For very thin metal skins (0.005-:J_nch stainless steel, e.g.) spot-welding the thermocouple on the inner surface of the skin is acceptable. The thermocouple leads should not be ,ecure4 on the outer surface· ,of· the model, since ice will anchor on the lead,&. If splicing of thermocouple leads is required, such splices should be made in a protected, constant-temperature location, outside the model in the test chamber. All thermocouple leads sh0'J.l.d be protected age.in�t moisture; asbesto9-covered wires are.not generally recommended for models in icing conditions. Shielded thermocouples are recommended for obtaining measurements of hot air temperature, although a trailing thermocouple such as that shown in figl...�e 50(b) is acceptable.
Tunnel Air Temp�ratl!I'e The twmel air tempera. ture is ob'ta:!.ned with a probe · t.bat separates the entrained water from the airstream as shown in figure 51. The probe consists of a nose _and rear cap and a �ousing containing a· temperature sensing element. Holes are located :i.n. the hQusing so that the a.tr flows.
thro'1gh the probe fran the rear to thi<: front of the probe. The'locations of these holes are based on pressu�e-distrib�tion studies. The water droplets, because of their inertia do Lot enter the housing at the rear ,.
locations. Because the nose-cap diameter is larger than the housing diam eter, the housing is protected. frcm icing. The nose cap is allowed to ice.· Thermal icing protection cm.,.lti. be incorporated in the nose cap; howe\re�, an error in the indicated tempera;ture would be incurred. The averll.ge iemperature-recovery factor for these probes is about 84'percent and is constant over a range of Mach numbers from 0.2 to 1.0. In the .
NACA icing tunnel the probes are used to mes.sure a:f:r temperature in the low-speed section upstream of the spray system and ahead of the contrac tion c�e. The air temperature measured by these probes is therefore essentially a total air temperature. Because the tunnel is always at - 18 - least saturated when the spray s:/stem is used, the ambient-air tempera ture-in the test section is computed by conventional wet-air equations (ref. 21). -The total air t�erature measured in the low-speed section of the tunnel is used for a base in these calculations.
Temperatures from Rotating Bodies o A typical means used to transmit the t.e.mperatures from a rotating _body to a recorder is shown schematically 1n figure 32. The thermocouple leads from the body are fed through a hollow motor shaft to the rear of a motor housing and through a thermocouple selector switch into a steam filled jacket 'that rotates with the shaft. Fram the rotating jacket, copper leads are attached to a slip-ring and brush assembly. Fram this assembly, copper leads are again led into a steam-filled stationary jacket. Copper-constantan leads are used from the stationary jacket to The steam jacket a thermocouple selector unit and to a flight recor,ier.
is used to provide a constant t�erature at critical junctions in the therm.ocouple·circuit, where the wi!'e metal in-the thermocouple leads is changed from copper and constantan to all copper and back again.
Pressure Tubes In icing conditions all pressure tubes subject to water impingement {pitot-static tubes, tubes in survey rakes and in the boundary layer, et� must be protected against icing. In the Lewis icing tunnel such pressure tubes are generally electrically heated.
A co-axial tube is used consist ing of two concentric tubes separated by a woven glass sleeving insulation (fig. 33). The ends of the tubes exposed to the airstree.m are silver soldered and shaped to obtain either static or total-pressure tubes. Tube sizes of 0.093- to 0.437-inch outside diameter and 0.057- to 0.393-inch inside diameter with wall thicknesses of 0.005 incl::. are in common usage.
These tubes are made of Inconei. A special tube bender was developed by NACA personnel to avoid collapsing of the co-axial tubing during bending of the tubes to a desired shape.
Surface Pressure from Rotating Body A scheme similar to that used to obtain temperatures from rotating bodies is also used to obtain surf'ace pressure measurements. Pressure lines from a model are fed into a hollow shaft (fig. 34). Each tube is then allowed to vent into a chamber composed of the hollow shaft, bear ings, and a stationary housing. All pressure sealing is accanplished by the bearings and felt seals. The pressure from each sealed chamber is then transmitted to a manometer. A water-jacket cooling system (not shown) is provided :for high-rotational-speed operation. This pressure system permits simultaneous readings of many pressures and is general.y limited only by the number of bearings used.
APPENDIX C
- 19 - APPENDIX C ICING INSTRUMENTS The-determination of liquid-water content and droplet-size dis tribution of natural and arti:ficial clouds has received considerable attention in connection with cloud physics studies and, in pertieula.r, in aircrat't icing studies. A knowledge o:f the liquid-water content and droplet-size distribution in clouds is o:f :fundamental importance in evaluating rate and area o:f ice :formation on various aircraft components, rate and area of erosion by impinging droplets on various aircraft sur :faces such as radomes, reduction af Visibility, attenuation o:f radar, and the basic mechanism of cloud :formation and precipitation.
Numerous methods for determining these pa.rametera have been proposed and tested, but each method suf:fers limitations as to accuracy or ease in obtaining or reducing tbe data to useful :form. In some cases the limita tions become very severe when measurements are attelllpted in high-speed a1rstreama. In addition to the rotating multicylinders and pressure type icing-rate meter discussed in the test, sane methods and techniques that have been widely employed to determine liquid-water content and/or droplet size are: (1) Cloud camera (2) Oil slides (3) Oil-stream aeroscope ( 4) Heated probes The methods a.re described and discussed in references 5.8 to 40.
The use o:f cameras (fig. 35) to photograph droplets directly in a cloud (ref. 58) i s based on f'undamen.tal ·principles and is ba.aically a sound technique, but there are practical di:f:ficulties. Because of the high magnif'ication required, the volume of the :field of View is extremely small. As a result, the average number of d:r-oplets per 8- by 10-inch picture is small in clouds of moderate liquid-water contents. Therefore, a large number of pictures are required in order to obtain a size distri bution. Since the magnification required is high, it i s difficult to design a camera so that the object plane is outside the undisturbed air stream about an airplane or camera mount.
The oiled-slide technique, where a glass slide covered with a suit- able oil is exposed to a droplet-J.atl,en airstream and is then photographed through a 1111-croscope, has been used all:lo to. determine droplet-size dis- .
tributions. This method yields photographs with a large number of droplets per picture from wbi-ch the droplet sizes may be measured. However, because of the relatively large size of the slide compared with the size of drop lets, the over-all collection efficiency and the iocal collection efficiencies - 20 - of tbe slide vary considerably with droplet size. The droplet-size dis tribution for any given area of the slide must be corrected according to the local collection efficiencies in order to obtain the true droplet size distribution of the cloud. The local collection efficiencies used for this correction a.re b9ised upon that of a ribbon in ideal two-dimensional flow. The exposure time required in order not to saturate the slide With droplets must be of the order of' a fraction of a second. This presents some difficulties, in that the slide must be moved rapidly or a protective cover must be opened and closed rapidly. This motion disturbs the air- flow field in the vicinity of the slide, and therefore the collection ef ficiencies of the slide are not the same as for a ribbon in ideal flow.
it.YJ. 0il-stream aeroscope -composed. of five main parts (fig. 36) - droplet pickup probe, circulating pumps for oil and air, photograp.'luc cell, light source, and a photomicrograpbic camera - :t:.as been developed. The droplet pickup probe cor-sists of a small-diameter tube w!th a small hole on one side. When operating, the probe is arranged so that the small ·hole faces upstream to the air flow carrying the cloud droplets. Oil is forced by a pump through the pickup probe in the direction indicated in the sketch. As the oil passes the small hole, any water droplets that enter are trapped in the oil. Oil does not flow O1..t of the droplet pickup hole, because the oil pressure is :na.inta:l..ned at atmospher:t� pressure by the air punp shown in figure 36. The oil containing the droplets then flows through the trans parent plastic cell where t�e droplets are photogJ;l'aphed with a photamicro graphic camera. The cha:-i.nel through the plastic cell narrows c.own at the point where the pictures are taken, so that all the cl.roplets are approximately + in the object plane of the camera. Af ,er lea.v:tng t:tle plastic cell, the oil passes throug:i a filt..::?r and trap where t.l:l.e water n.roplets are r.emoved. The droplet size and distribution can then be f"'.et.ermJned. by measur:J.ng the images on the phctcgraphs from the known magnifics.tior:i.. After tee droplet distri bution is known, the liquid-water content of' the cloud can be calculated from the known geometry of the inatrttment, the airspeed, anc the oil-flow rate.
Limited data indicate that this instrumer::t &hows excellr-:mt promise for ob taining the desired informb.tion.
The heate:i-wire instrument ccr.sists bn.sicall.v of a lO<Jp of resistance wire (refs. 39 and 40) which is :nou.r:.ted. :tn t'ht; al.rstream (fig. 37) and is heated electricaJ_ly by passin� currP-Jnt +�b.z·u!.lg!-' the wir�. The wiri::: diameter is 0.021 to 0.054 inch, with a maximum po�Jr input of 31 to 300 watts, respectively (ref. 40). The change in wire r'=sists.nce from the clear-air condition, resulting from cooling d,1e to zvap:.ration of impinging cloud water droplets, is used as a meast1re cJf th,e liquid-water content, or icing severity. Although the hea.te·5.-wire instrument hs.s seY�ral disadvantages as pointed out in reference 40, a. workable instrument can be obtained that is very useful in studying cloud m:!.crostructure.
A variat:!.on of the heated-wire instrumer-.+-, is �urrcntly under development at the NACA Lewis laboratory. T.r.:.is instrunYm�s cons. a�;s of a he.ated tube - 21 - operating at constant surface temperatt:7e With a variable power input.
�s instrument has the f'oll,owing advantages over a constant-power hea.ted wire probe: (1) The measured cha.nge in required power is linear with water impinge ment, becau� the surf'ace temperature is f'ixed.
(2) The change of heat-tran&f'er coef'f'icient under all conditions is · minimized.
(3) The sensitivity to water impinge111ent is ma.x:l.mized.
(4) The power input is easily :measured.
(5) With proper design, the time c·onstaI?.t is less that that• of' a wire heated with constant power.
v While the control circuit for this probe has pro -ed f'or.nida.ble, a satis factory unit has been designed and operated. The probe is currently being calibrated ,in t..lie Lewis icing tunnel.
- 22 - REFERENCES 1. Ha.cker, Paul T., and Dorsch, Robert G.: A Summary of Meteorological Conditions Associated with Aircraft Icing and a Proposed Method of Selecting Design Criteriona for Ice-Protection Equipnent. NACA TN 2569, 1951.
2. Lewis, William, and Bergrun, Norman R.: A Probability Analysis of the Met·eorological Factors Conducive to Aircraft Icing in the United States. NACA TN 2738, 1952.
3. Jones, Alun R., and Lewis, William: Recormnended Values of Meteorologi cal Factors to be Considered in the Design of Aircraft Ice-Prevention Equipment. NACA TN 1855, 1949.
4. Lewis, William, Perkins, Porter J. , and Brun, Rinaldo J.: Proced,.:..re for Measuring Liquid-Water Content and Droplet Sizes in Super- cooled Clouds by Rotating Multicylinder Method. NACA RM E53D23, 1953.
5. Perk.ins, Porter J., McCullough, Stuart, and Lewis, Ralph D.: A Sim plified Instrumen t for Recording and Indicating Frequency and Inten sity of Icing Condit:1.ons Encountered. in Flight. NACA RM E51El6, 1951.
6. Brun, Rina.ido J., Gallagher, Helen M., and Vogt, Dorothea E.: Impinge ment of Water Droplets on NACA 65A004 Airfoil and Effect of Change in ° Airfoil Thickness from 12 to 4 Percent at 4 Angle of Attack. NACA TN 3047, 1953.
7. Brun, Rinaldo J., Gallagher, Helen M., and Vogt, Dorothea E-.: Impinge ° ment of Water Droplets on N.ACA 65 _-208 and 65 -212 Airfoils at 4 1 1 Angl.e of Attack. NACA TN 295.2, 1953.
8. Guibert, A. G., Janssen, E., .and Ro'J)bins, W. M.: I>etermination of Rate, Area, and Distribution of Impi!lgement of Waterdrops on Various Airfoils from Trajectories Obtained on the Diffe:rent:1.al Analyzer. NACA RM 9A05, 1949.
9. Brun, Rinaldo J., Gallagher, Helen M., and Vogt, Dorothea E.: Impinge ment of Water Droplets on NACA 65A004 Airfoil at a Angle of Attack.
NACA 'l'N 3155, 1854.
· 10. Brun, Rinaldo J., Serafini, John So, and Gall.a.gb.er, Helen M.: Impinge ment of Cloud Droplets on Aerodynamic Bodies �� Affected by Compressi bill ty of Air Flow Around the Body. NACA TN 2903, 1953.
11. Dorsch, Robert G., and Brun, Rinaldo J.: A Method for Determining Cloud-Droplet Impingement on Swept Wings. NACA TN 2931, 1953.
- 2:5 - 12. Brun, RlnaJ.do J., and Dorsch, Robert G.: Impingement of Water Droplets on an Ellipsoid w1 th Fineness Ratio 10 in Axis�tric Flow. NACA TN 5147, 1954.
1:5. Dorsch, Robert G., Brun, Rlnaldo J., and Gregg, John L.: Impingement of Water Droplets on an Ellipsoid with Fineness Ratio 5 in .Axisymmetric Flow. NACA TN 5099, 1954.
14. Brun, Rinaldo J., and Mergler, Ha.rry W.: Impingement of Water Droplets on a Cylinder in an Incompressible Flow Field and Evaluation of Rotating Multicylinder Method for �asurement of Droplet-Size Dis tribution, Volume-�dian Droplet Size, and Liquid-Water Content in Clouds. NACA TN 2904, 1955.
15. von Glahn, Uwe !I., Gelder, Thomas F., and Smyers, William H., Jr.: A Dye-Tracer Technique for Experime:itally Obtaining Impingement Char acteristics of Arbitra.ry Bodies and a Method for Determining Droplet Size Distribution.
NACA TN 555.8, 1955 • 16. Sherman, P., IQ.ein, J. S., and Tribue, M.: Determination of Drop Tra jectories by Means of an Extension of Stokes' Law. Eng. Res. Inst., Univ. Mich., Apr. 1952. (Air Res. b.r..i DeY. Command, US.AF, Contract AF 18(600)-51, Proj. M992-�.)
17. von Glahn, Uwe H., and Gray, Vernon H.: Effect of Ice Formations on Section Drag of Swept NACA 63A-,J09 .Airfoil. w1 th Partial-Span I.P.a.ding Ed.ge Slat for Various Modes of Thermal Ice Protection. NACA .RM E.55J50, 1954.
Effect of Ice and F'rost Forma 18. Gray, Vernon H., and von Gla!ln, Uwe li.: tions on Drag of NACA 65 -212 Airfcil for Various M::>des of Th'=!r.na.1 Ice Protection. NACA TN 2962, 1955.
19. Lewis, James P., and Blade, Rob":):rt J.: Bxperime�tal Investigation of Radome Icing an� Icing Protection. NACA TM E52..T51, 1953.
20. Nee1, Carr B., Jr. , BergrJ.n, Norma.n R. , J·:.:.k.::>f:f, Da."'l"'id., a.n,i s�Uaff, Bernard A. : The Calculation of the !feat Required for Wing Thermal · Ice Prevention in Specified Il.!i!'ig Cor..di-l;ions. NACA TN 1472, 1947.
21. Gelder, Thomas F., and lewis, James P.: Comparison of Heat Transfer from Airfoil in Natural and Simulate� Icing Conditions. NACA TN 2480, 1951.
22. von Glahn, U.: Preliminary Results of Heat Transfer from a stationary and Rotating Ellipsoidal Spinner. NACA RM E55F02, 1955 • 25. Gelder, Thomas F., Iewis, James P., an.d. Koutz, Stanley L.: Icing Protec tion for a Turbojet Transport Airplane: Beating Requirements, Methods of Protection, and Performance Penalties. NACA TN 2866, 1955.
- 24 - 24. Hardy, J. K. : Kinetic 'l'emperature of Wet Surfaces - A Method of Cal-culat:1Jp.g the Amount of Alcohol Required to Prevent Ice, and the Derivation of the Psychrom-stric Equation. NACA WR A-8, 1945.
(Supersedes NACA ARR 5Gl3.)
25. Hardy, J. K.: An Analysis of the Dissipation of' Feat in Conditions of Ic ing from a Section of the Wing of the C-46 Airplane. NACA Rep.
831, 1945. (Supersedes NACA ARR 4Illa.)
26. Gray, V. H., Bowden, D. T., and von Glahn, U.: Preliminary Results crf Cyclical De-Icing of a Gas-Heated Airfoil. NACA. RM E51J29, 195 2.
27. Gray, Vernon H., and. Bowden, Dean T.: Comparison of Several Methods f of Cyclic De-Icing of a Gas-Heated Air oil. NACA RM E5:3C27, 1953.
28. lewis, James P., and Bowden, Dean T.: Preli!ll:1.nary Investigatioz:. of Cyclic :::)e:-Icing .bf an Air:fc-11 Using an External Electric l!eater.
NACA RM E51J30, 195 2 •.
29. Coles, Willard D.,. and Ruggeri, Robert S. : Experimental Investigation of Sublimation of Ice at Subsonic and Supersonic Speeds and Its Rela.- tion to Heat Transfer. NACA 'l'N :3104, 1954:.
:50. Coles, Willard, Rollin, Vern G., and Mt..ll.b.olland, Donald R.: Icing Protection Req�rements for Reciprocating-Engt�e Ir.duction Systems.
KA.CA Rep. 982, 1950. (Supersedes NACA. TrT 1993.)
:51. Acker, Lo�n W.: Natural Icing of an Axial-Flaw Turbojet Engine in Flight i'or a Single Icing Conditiot1. NACA RM E8F01a, 1948.
32. Acker, Loren W.: Preliminary Rl:lsults of N'a.t-:.iral Icing of a,n Axial-F'low 'furbojet Engine. NACA RM E8Cl8, 1948.
33. Gray, Vernon H., and Bowden, Dean T.: Icing Characteristics and Anti Icing Heat Requirements for Hollow ar.t: �\.:terr...ally M:>dified. Gas-Jleq,-ted ' Inlet Guide Vanes. NA.CA BM ESC,108, 1950.
34. Callaghan, Edmund E., and �ra.f'ir.d, Jo:::ir,!, S.: A Met:toa. for Rap:1.d ..Deter mination of the Icing Li.mit of a Body in 'I-erms of the Stream Condi tions. NACA TN 2914, 1953.
35. Callaghan, Edmund E., and Serafini, Jor.n S.; Analyt1ca.l In,;estigation of Icing Limit f'or Diamond-Shaped Airfoil in Tre,nsonic and Supersor.ic Flow. NACA TN 2861, 1953.
:56. Coles, Willard D.: Icing Limit and Wet-Surface Temperature Variation for Two Airfoil �pes under Simulated High-Speed Flight Conditions.
NACA TN 3396, 1955.
- 25 - ' 37. La�r, Irving, and Blodgett, Katherine B.: A Mathematical Investigation ,· · of Water Droplet Trajectories.
Tech. Rep. No. 5418, Air Materiel Command, :, AM', Feb. 19, 1946. (Contract No. W-33-038-ac-9151 with General Electric eq .. ) 38. McCullough, Stuart, and Perkins, Port.er J.: Flight Camera for Photo graphing Cloud Droplets in Natural Suspension in the Atmosphere. NACA . RM E50K0la, 1951.
39. Neel, Carr B. , Jr. , and Steinmetz, Charles P. : The Calculated and Measured.
Performance Characteristics of' a Heated-Wire Liquid-Water-Content Meter ' for Measuring Icing- Severity. NA.CA TN 2615, 1952.
40. Neel, Carr B.: A Heated-Wire L:1.g;uid-Water Content Instrument and, :Results of' Initial Flight ��st in Icing Conditions. !lACA RM A54I23, 1955 • SEVERE ICING OF AIRCRAFT COMPONENTS COMPRESSOR INLET IC1NG Figure 1.
AIRCRAFT SURFACES REQUIRING ICING PROTECTION
"-,
� "' \ /
"
' /
"' \
"'/
SHADING INDICATES AREAS SUBJECT TO ICING Figure 2o
DROPLET TRAJECTORIES ABOUT AIRFOIL
AIR STREAMLINES
!CS-10718 !
DROPLET PATHS Figure 3.
Figure 4.
PRESSURE-TYPE ICING RATE METER SUPPORT TUBES TOTAL-PRESSURE HOLES IN LEADING EDGE OF PROBE AIR FLOW TRANSFORMER TO SUPPLY HEATING CURRENT TO PROBE A. C. POWER SOURCE Figure 5.
WORLD ICING SURVEY ROUTES
• r r , r ....
•
0 C' ••• . ) .
jcs-1011s j Figure 6.
WATER-DROPLET -TRAJECTORY ANALOG
Figure 7 •
LOCAL IMPINGEMENT RATES
ON A 15-PERCENT THICK
SYMMETRICAL AIRFOIL
° ANGLE OF ATTA CK, 0 ; WATER CONTENT, 0.7 GM/CU. M.
AIRSPEED , 150 KNOTS; CHORD, 8 FT I.LI � 0:: 1- z I.LI
�t 6
�� a. �
:! I
0::' 4 m I.LI I- ..J ; � ES-10816J 0 .01 .02 .03 .04 .05 .06 .07 .oe .09 SURFACE DISTANCE MEASURED FROM LEA DING EDGE/CHORD Figure a.
Figure 9.
COMPARISON OF THEORETICAL AND EXPERIMENTAL IMPINGEMENT DATA
NOTE: IMPINGEMENT PARAMETER BASED o EXPERIMENT AL ON HEAN DROPLET SIZE I BY VOLUME I THEORETICAL I- z "' z ��30 NOTE: IMPINGEMENT PARAMETER
�t
BASED ON MAXIMUM DROPLET SIZE '1!': ci D-o:: IN DROPLET SIZE DISTRIBUTION
� §
o, I- l&J zV l&JZ ,o
E�
O 2:
i�
� ---------- �; --�--���� .... - � ...._ __._____._____............................
1 .0 . 1 1 . 1 0 0 oo 2: � IMPINGEMENT PARAMETER, flDROP SIZE , AIRSPEED. COMPONENT SIZE I Figure 10.
TYPICAL AIRFOIL ICE FORMATIONS
t
t
�
@!:Tu@
(8) DOUBLE PEAK GLAZE ICE.
(A) RIME ICE. DATUM
DATUM AIR TEMPERATURE,
AIR TEMPERATURE, 00 F
°
50 F. HIGH RATE OF WATER CATCH
Figure 11.
PLAN VIEW OF ICING TUNNEL
VENTILATING TOWERS HEAT EXCHANGER REFRIGERATION COILS SPRAY SYSTEM Figure l2o ICING TUNNEL SPRAY SYSTEM Figure 15.
MODEL USED TO STUDY AERODYMAMIC PENALTIES CAUSED BY ICING Figure 14.
•
CHANGES IN LIFT AND DRAG CAUSED BY ICING
INITIAL LIFT COEFFICIENT, 0.20 INITIAL DRAG CO[FFICIENT, ifj.008 0.0068 u_oo u ICING TIME, MIN
m
Figure 15.
TYPICAL HEAT LOSSES ASSOCIATED WITH ANTI-ICING
·4000 HEAT TO IMPINGING WATER / TOTAL !£AT fl.OW g r:L' ::c - EVAPORATIVE HEAT LOSS ' 2400 � AREA OF IMPINGEMENT � 1600 \ � /
' /
--
�� --
CONVECTIVE HEAT LOSS
'
!CS-10774 !
4 8 24 0 12 16 20 DISTANCE ALONG SURFACE/CHORD, PERCENT Figure 16.
C·46 AIRCRAFT WITH
THERMALLY ANTI-ICED AIRFOIL MODEL
Figure 17.
TYPICAL ELECTRIC HEATER CONSTRUCTION FOR C-46 FLIGHT STUDIES • THERMOCOUPLE PLANES 1/52" 0.006-IN. ALUMINUM SKIN PLASTIC IMPREGNATED FABRIC, 1/64-IN. THICK ELECTRICAL RESISTANCE HEATER STRIP 1/2 BY 0.002 IN.
Figure 18.
♦ NACA AIR HEATED CYCLICALLY DE-ICING WING l VALVE PORT (PLENUM CHAMBER INLET) REAR PASSAGE CHOROWISE GAS FRONT SPAR FLOW PASSAGES PARTING STRIP SECTION A-A CENTER SEGMENT Figure 190
ICE FORMATIONS ON AIRFOIL BEFORE AND AFTER
CYCLIC HEATING
'
!CS-10765 !
18) AFTER IS-SECOND HEAT-ON PERIOD
(A) BEFORE HEAT-ON PERIOD
Figure 20.
PNEUMATIC BOOT DE-ICING SYSTEM
Figure 21 ICING PROTECTION FOR PISTON ENGINE-INDUCTION SYSTEM SPEED. DENSITY METERING, WATER-EXCLUDING AND CYLMDER-FUEL INJECTION SUBMERGED INLET ELBOW Figure 22.
TYPJCAL ICING OF JET ENGINE INLET
...
- INLET GUIDE VANE INLET UP ,,,.- ISLAND \· ACCESSORY HOUSING STATOR BLADES SECOND STAGE FIRST STAGE SECOND STAGE ICS-10764!
Figure 23.
MODEL USED IN GUIDE VANE STUDIES
'
Figure 24.
HEAT SAVINGS OBTAINED BY GUIDE VANE PARTITIONING
TYPICAL HEAT FLOW, BTU/HR FOR SEVERE IONG CONDITION ,
C-------
(Al VANE I. FULLY HOLLOW (Bl VANE 2. INTERNAL FIN AND INSET r- , lNSERT (Cl VANE 3. INTERNAL INSULATING INSERT ics-101111 Figure 250 LEWIS ICING TUNNEL FACILITY ,r W'NTILATING TOWERS - TEll'ORARY TEST SECTION FOR PAOP£LLER ICING IIESEAll0i IAI AERIAL - Of ICNG 1\NEL CBI PHANTOM ORAWNG Of LEWIS ICING 1\NEL AHO OFFICE IIUl.DIIG � Figure 26.
CROSS SECTION OF NACA AIR-WATER ATOMIZING SPRAY NOZZLE
SPRA V STRUT FAIRING J:5-lOBllJ Figure Z,.
NACA 3.84x 10-INCH HIGH-SPEED ICING DUCT TUNNEL
REFRIGERATED AIR SUPPLY ....
PLENUM CHAMBER THERMOCOUPLE PLANE FLOW-STRAIGHTENING TUBES 11• DIA.I
�. f 12·
k--.._--_ -_-_:::-_::-........Jy
�SCREENS I
----l--+-----120"----1 �
'
BELLMOUTH ENTRY Figure 28.
NACA 12 x 12-INCH HIGH-SPEED ICING DUCT TUNNEL
ALTITUDE SPRAY EXHAUST NOZZLES
I
I
STRAIGHTENING
t
VARIABLE VAN ES REFRIGERATED STEAM SECOND HEATED AIR THROAT THROAT Figure 29.
THERMOCOUPLE INSTALLATION DETAILS NO. 32 AWG GLASS ENAMEL PUSHBACK WIRE OUTER COVERING .01� TO GLASS WRAP .020 DIA. � - ENAMEL _; PEENING TOOL SKIN PEEN HERE AT 6 POINTS __},.t/ .025 TO .030 DIA.
AROUND PERIPHERY -4j -- TO FIT T.C. BALL 1/8 OIA.Q IAI SURFACE THERMOCOUPLE NO. 32 AWG AFTER SILVER 80L.Df:RING GAS FLOW .....
� AFTER HONING 0.06 O.D. X 0.006 WALL _J CERAMIC TUBING STAINLESS STEEL TUBING JUNCTION DETAILS IBI AIR FLOW THERMOCOUPLE Figure 50.
INERTIA-SEPARATION TEMPERATURE PROBE
5/32 DRILL MICARTA INSERT MOUNTING TUBE THERMOCOUPLE LE ADS ps-1oaOSJ losE CAP 10 HOLES, 5/32 DR ILL I THERMOCOUPLE JUNCTION
I I [
ENCASED IN SCALE, IN.
STAINLESS-STEEL TUBE Figure 31
SCHEMATIC DIAGRAM FOR TEMPERATURE MEASURING SYSTEM
ON ROTATING BODIES
DISCONNECT PLUGS SLP-RNG ASSEMBLY BRUSH \/ THERMOCOUPLE SWITCHING UNIT \ Cu ::::---� C u TO THERMOCOUPLE } - SELECTOR BOX AND ,---- - FLIGHT RECORDER C C C ROTATING STEAM JACKET STATIONARY STEAM JACKET Figu!e 32
DETAILS OF ELECTRICALLY HEATED PRESSURE TUBES
-- - -----f+ DETAIL OF DETAIL OF TOTAL PRESSURE TUBE STATIC PRESSURE TUBE TYPICAL ELECTRICALLY HEATED RAKE Figure 33
SCHEMATIC DIAGRAM FOR PRESSURE MEASURING SYSTEM
ON ROTATING BODIES
HOLLOW STATIONARY HOUSING FELT SEAL AND BEARINGS
�� r-----,11llllla1llllla��il'llllJZ--l71====
ROTATING LEADS TO MODEL PRESSURE TAPS PRESSURE LEADS TO MANOMETER HOLLOW ROTA TING SHAFT ts-100151 - - DIRECTION OF TRANSMISSION Figure 34
NACA CLOUD CAMERA
'1
--------------------------r
---- ROTATING MIRROR �
\,.,,i
,
--��--i---- J' ! FILM
_j J � ROJECTION LENS P I I I I � � +------- : : ----�----------- i I i I -- --- ___ ----L.. _________ .,, J I I I I I Y- AIR LANE SKJN I P I I \ I OBJECTIVE LENS
i !
\ I I I, � ',. ) FIRST IMAGE OF LIGHT SOURCE
lBJEr:r
PLANE URCE fs-i�O �� -������!;��) .... _______ .,, Figure 35
SCHEMATIC SKETCH OF OIL-STREAM AEROSCOPE
AIRFOIL SHAPED DIRECTION OF n1---PICKUP PROBE, 0.156" I 0.50" DROPLET MOTION/ MOUNTING STRUT .04" Oil. OROPLET ,....__...,, P ICKUP HOLE AIR CRAF T SKI N AIR BUBB LES IN OIL - DIRECTION OF OIL Flow----.
� CLOUD DROP LE TS IN •0,1 TRANSPARENT PLASTIC THREE·WAY STOPCOCK ' PHOTO FL ASH LAMP CHANNE L DIMENSION
. ooe· IN DIRE CTION OF
OPTICA L AXIS. A ND t!'l-lC8(1/1 .1" lT RIGHT ANGLE S FILTER ANO WATER TRA P OIL RESERVOIR Figure 36 HOT WIRE LIQUID WATER CONTENT METER (Al WIRE-LOOP ANO SUPPORTING-STRUT ASSEMBLY WIRE LOOP WIRE DIAMETER, 0,021 IN, EXPOSED LENGTH, 1,63 IN, � (Bl CONSTRUCTION DETAILS OF WIRE-LOOP ASSEMBLY ANO SUPPORTING-STRUT ASSEMBLY Figure 37